SPACE CADRE - SPACE COMMAND - SPACE FORCE

We dedicate this work to all four
thousand million year-old bacteria
who may once have lived in our Solar
System, and to all life everywhere
created, sustained, and sometimes
destroyed by suns. With all our hopes
for the future,

The Ra Team

Final

Report

^he S un
jot
Science

International Spac&t-jffive^ |

© 1996 International Space University, All Rights Reserved

The 1996 Summer Session of the International
Space University existed for ten weeks at the
Technical University of Vienna, hosted by the
Austrian Society for Aerospace Medicine.

The cover image of the Sun was taken by the Solar and Heliospheric
Observatory Extreme Ultraviolet Telescope. The wavelength
shown is 195 Angstroms, revealing highly ionised iron atoms in the
lower corona at 1.5 million Kelvin. The North and South poles of
the Sun clearly show coronal holes, a phenomenon not yet fully
understood. The image was courtesy of the SOHO EIT Consortium
(SOHO) is a joint endeavour by ESA and NASA.

Additional copies of the Design Project Executive Summary or the Full Report , Ra: The Sun for Science and
Humanity, may be ordered through the ISU Headquarters in Strasbourg or the ISU North American Office .

International Space University

Strasbourg Central Campus

Parc dTnnovation

Boulevard Gonthier d'Andemach

67400 ILLKIRCH-GRAFFENSTADEN

FRANCE

Tel: +33 (0)3 88 65 54 30
Fax : + 33 (0)3 88 65 54 47

International Space University
North American Office
3400 International Drive, NW
Suite 4M-400

Washington, DC 20008-3098
USA

Tel: +1 (800) 6771987 (USA and Canada only)
Tel: +1 (202) 2371987 (other countries)

Fax: +1 (202) 237 8336

See the ISU website at http://www.isunet.edu/

"I would say that man should live for loving, for
understanding, and for creating. I think man should spend all
his ability and all his strength on pursuing all these three
aims, and he should sacrifice himself, if necessary, for the
sake of achieving them. Anything worthwhile may demand
self-sacrifice, and, if you think it worthwhile, you will be
prepared to make the sacrifice."

Arnold Toynbee, Surviving the Future
Oxford University Press, 1971.

Over this summer at ISU, we spent most of our abilities and
our strengths on appreciating each other and on
understanding what "the Sun for Science and Humanity"
could mean. Sacrifices have sometimes been necessary, and it
was worthwhile. Here is what we have created...

The Ra Team.

Ra: The Sun for Science and Humanity

Acknowledgements

To everyone who contributed their time, energy, and expertise to the Ra team, we
wish to express a heartfelt: Danke, Thank you, Merci, Grazie,"H'^H <: li<s , Gracias,
Gracies, Spasibo, qJapeODwYc, Takk, Bedankt, Tack, Kiitas, Cam-on, Tak!

Sponsors

ESA

NASA Office of Policy and Plans
Bradford Engineering

Design Project Team

Burke, James D.
Scoon, George
Lamontagne, Chantal
Mallory, Gregory

ISU Team

Caltech Jet Propulsion Laboratory, USA
European Space Agency, ESTEC , UK
Carleton University, Canada
University of New Brunswick, Canada

Atkov, Oleg
Becker, Francois
Bousquet, Michel
Cohendet, Patrick
Crosby, Norma
De Dalmau, Juan
Duchesne, Simon
Fazio, Giovanni
Green, James
Kendall, Ehvid
Johnson-Freese, Joan
Hamilton, Douglas R.
Logsdon, John
Marov, Michail
Mironjuk Nadezhda
Mendell, Wendell
Pieson, Dmitry
Pelton, Joseph
Rycroft, Michael
Sanders, Berry

Research Cardiovascular Centre, Russia
International Space University, France
International Space University, France
Universite Louis Pasteur, France
Observatoire de Meudon, Denmark
European Space Agency, Spain
Canadian Armed Forces, Canada
International Space University, USA
Goddard Space Flight Center, USA
Canadian Space Agency, Canada
Air University, Maxwell AFB, USA
University of Calgary, Canada
George Washington University, USA
Russian Academia of Science, Russia
Moscow Aviation Institute, Russia
NASA Johnson Space Center, USA
Moscow Aviation Institute, Russia
International Space University, USA
International Space University, UK
Bradford Engineering, The Netherlands

European Space Agency, France

Tuinder, Paul, H.

Visiting Lecturers

French, Lloyd C. Caltech Jet Propulsion Laboratory, USA

Nakatani, Ichiro Institute of Space and Astronomical Science,

Japan

Randolph, James E. Caltech Jet Propulsion Laboratory, USA

Vaisberg, Oleg Institute for Cosmic Research, Russia

Weinstein, Stacy S. Caltech Jet Propulsion Laboratory, USA

Willikens, Philippe European Space Agency, Belgium

Worden, Pete S. HQ Air Force Space Command, USA

Wnuk, Gerard European Space Agency, ESTEC, France

Correspondence and Teleconference

Ankersen, Finn
Asmar, Sami
Atkers, Lisa D.
Andersson, Mats
Bainum, Greg C.

Balts, Keith W.
Bourke, Roger D.
Bravo, Sylvia
Buttighoffer, Anne
Bolduc, Leonard
Carrington, Connie K.
Dale, Gary E.
Dawson, Simon
Detman, Thomas R.
Emslie, Gordon
Flowers, Nick
Friedman, Louis
Fujita, Toshio
Gruenagel, R.

Hajos, Gregory
Harett, G Shaw
Hilgers, Alain
Huber, Martin
Huber, Ralf
Jackson, Bernard V.
Juhlin, Lars-Erik
Kakuda, Roy Y.
Kahler, Steve
Keil, Steve, L.
Lanzerotti, Louis J.

European Space Agency, ESTEC, Holland
Caltech Jet Propulsion Laboratory, USA
50 th Weather Squadron, USA
Sydkraft, Sweden

Air Force Institute of Technology, USA
1 st Space Operation Squadron, USA
Caltech Jet Propulsion Laboratory, USA
Institute de Geofisica, Mexico
Observatoire de Meudon, France
Hydro Quebec, Canada
NASA, MSFC, USA

California Institute of Technology, USA

Microcosm, USA

Space Environment Center, USA

University of Alabama, USA

Mullard Space Science Laboratory, UK

Planetary Society, USA

Caltech Jet Propulsion Laboratory, USA

European Space Agency, ESTEC, Holland

NASA, MSFC, USA

Institute for Australasian Studies, Australia
European Space Agency, ESTEC, Holland
European Space Agency, ESTEC, Holland
DLR, Germany

University of California, San Diego, USA

ABB Power, Sweden

Caltech Jet Propulsion Laboratory, USA

Phillips Laboratory, USA

National Space Observatory, USA

AT&T Bell Laboratories, USA

Lesh, James

Caltech Jet Propulsion Laboratory, USA

Lindsay, Gretchen

50 th Weather Squadron, USA

Leibacher, John

GONG, USA

Lundstedt, Henrik

Lunds Universitet, Sweden

Marsden, Richard

European Space Agency, ESTEC, Holland

Martin, Tony

AEA Technology, UK

McCray, Joel D.

50 th Weather Squadron, USA

Morabito, David

Caltech Jet Propulsion Laboratory, USA

Mulqueen, Jack

NASA, MSFC, USA

Neugebauer, Marcia

Caltech Jet Propulsion Laboratory, USA

Noca, Muriel

Caltech Jet Propulsion Laboratory, USA

Oberger, Kjell

Elforsk, Sweden

Oppenhauser, G.

European Space Agency, ESTEC, Holland

Perkinson, Don T.

NASA, MSFC, USA

Price, David

NASA, MSFC, USA

Rahe, Juergen

NASA, Office of Space Science, USA

Sauer, Carl G.

Caltech Jet Propulsion Laboratory, USA

Saunders, Stephen R.

NASA, Office of Space Science, USA

Schlingloff, Hanfried

Ingenieurburo "Dr Schlingloff", Germany

Scro, Kevin D.

50 th Weather Squadron, USA

Seitz, David B.

1 st Space Operations Squadron, USA

Sercel, Joel C.

Caltech Jet Propulsion Laboratory, USA

Singer, Howard

Space Environment Center, USA

Stem, Bob

Lockheed Martin, USA

Tarvin, Christina A.

50 th Weather Squadron, USA

Taur, Roger R.

China

Lockheed Martin, Harbin Institute of Tech.,

Thompson, Richard

IPS Radio and Space services, Australia

Tschan, Christopher R.

50 th Weather Squadron, USA

Tsurutani, B. T.

Caltech Jet Propulsion Laboratory, USA

Verbaas, Ad

Fokker Space BV, The Netherlands

Yuen, Joe

Caltech Jet Propulsion Laboratory, USA

Zwickl, Ron

Space Environment Center, USA

• Ra: The Sun for Science and Humanity

Authors

Marc ABELA

Canada

Development Engineer

Canadian Space Agency - University of Tokyo

Andrew J. BALL

United Kingdom

Ph.D. Student

University of Kent

Christopher BARRINGTON-LEIGH

Canada

Graduate Student in Applied Physics

Stanford University

Soren Abildsten B0GH

Denmark

Ph.D.Student, Department of Electronic Engineering,
Aalborg University, Denmark

John BUCKLEY

United Kingdom

Higher Scientific Officer, Space Department

Defence Research Agency, United Kingdom

Frank BUDNIK

Germany

Research Scientist,

Institute for Geophysics and Meteorology,
Braunschweig, Germany

Jean-Pierre (J-P) CACHELET

France

Ph.D in Aerospace Design and Optimization

College of Aeronautics, Cranfield, UK

Giovanni CARRA

Italy

MS in Electronic Engineering

Directorate of Launchers

European Space Agency, Paris

Guillem CHUST

Spain

Biologist, Remote Sensing and Ecology

University of Barcelona, Spain

Juan M. del CURA

Spain

Aerospace Engineer, Aerospace Division, SENER
Associate Professor, Aerospace Department,
Polytechnic University, Madrid

Vincent DELHAES

France

Engineer

Military and Space Technology

AEROSPATIALE, France

Robin FLACHBART

USA

Aerospace Engineer

NASA, Marshall Space Flight Center

Alexander GICZY

USA

Captain, United States Air Force

Onizuka Air Station, California

Mariella GRAZIANO

Italy

Graduate Student in Aerospace Engineering
University La Sapienza, Rome, Italy

John HASSE

USA

Graduate Student of Geography

Rutgers University, New Jersey

Hannu HOLMA

Finland

Graduate Student in Space Physics

University of Oulu

Maxim V. JACOBSON

Russia

Graduate Student of Aerospace Engineering

Moscow Aviation Institute, Russia

Xiaoguang JIA

China

Professor, School of Astronautics

Harbin Institute of Technology, Harbin, China

Jean-Yves JOUAS

France

R&D Project Engineer

Societe Europeenne de Propulsion (SEP)

Ateet KAPUR

France

Young Graduate Trainee

European Space Operation Center, Darmstadt

Yasuharu KAWABATA

Japan

Graduate Student in Aerospace Engineering

Tohoku University, Sendai, Japan

Stefan KOGL

Germany

Structural Engineer

Oerlikon-Contraves AG, Zurich, Switzerland

Gunther LIENTSCHNIG

Austria

Graduate Student in Physics

Vienna University of Technology

Xavier LOBAO

Spain

Telecommunications engineer

Project Manager, On-Board Data Handling

Indra Espacio, Barcelona

Alexandre MARTYUSHOV

Russia

Student in Aerospace Design

Moscow Aviation Institute

Olaf MASTENBROEK

The Netherlands

Research Engineer

National Aerospace Laboratory NLR, The Netherlands

Jeffrey T. MORISETTE

USA

Ph.D.Candidate in Forestry / Remote Sensing

North Carolina State University

Cuong Q. NGUYEN

USA

Aerospace Engineer

NASA, Johnson Space Center

Ena L. NISHIMUTA

USA

Aerospace Engineer

NASA, Marshall Space Flight Center

Yamal Chandra RAJBHANDARY

Nepal

Graduate Student

The University of Iowa, USA

Hermen M. REHORST

The Netherlands

Aerospace Engineer

Delft, The Netherlands

Daniel A. REY

Canada

Ph.D. Candidate, Mechanical Engineering

McGill University, Montreal Canada

Wolfgang REINPRECHT

Austria

Research Scientist

Technical University Graz, Austria

Stephanie A. ROY

USA

Graduate Student in Science, Technology, and Space
Policy

Space Policy Institute, George Washington University

Rie SAKAKIBARA

Japan

NASD A

Tsukuba Space Center

Robie SAMANTA ROY

USA

Research Staff Member

Institute for Defense Analyses, Alexandria, VA

Francesco SARTI

Italy

Electrical Engineer

Attitude & Orbit Control Systems, ESA/ESTEC

Martin SILL£N

Sweden

Graduate Student in Information Technology

Royal Institute of Technology, Stockholm

Kristofer SKAUG

Norway

Human Being of Norwegian Origin

M.Sc. Delft University of Technology

Isabella C. SKRNA-JAKL

Austria

Ph.D in Mechanical Engineering

Institute of Lightweight Structures and Aerospace
Engineering, Vienna University of Technology

Brant L. SPONBERG

USA

Space Policy Graduate Student

The George Washington University, Washington, DC

Martin TAJMAR

Austria

Graduate Students in Physics

Vienna University of Technology

Randal TEDROW

USA

Captain, United States Air Force

Air Force Space Command, Colorado

Jean Daniel (JD) TESTS

France

Commandant, Armee de l’Air

Etat Major des Armees, Bureau Espace

Centre National d’Etudes Spatiales, Toulouse, France

Cornelia THIEME

Germany

Aerospace Engineer

MAN Technology, Munich

Yumi TOMITA

Japan

Graduate Student in Aerospace Engineering

Nihon University, Japan

Marianne Kronstad VINJE

Norway

MSc in Informatics / Mathematical Modelling
Researcher, Kongsberg Aerospace

Zemin WANG

China

Senior engineer

Xichang Satellite Launch Center

Mathias WUHR

Canada

M.Eng. graduate in Aerospace Engineering

Carleton University, Ottawa, Canada

Yanjun XU

China

Senior Engineer

Taiyuan Satellite Launch Center (TSLC)

Huiqin YANG

China

Senior Engineer - Luoyang Institute Tracking and
Telecommunications Technology, China Satellite
Launch Tracking and Control General, China

Olga ZHDANOVICH

Russia

Ph.D.Candidate Remote Sensing / GIS

Russian Academy of Sciences, Moscow

Xian zheng ZHU

China

Director of CEE Center

China Committee of Sci-Tec & Industry for National
Defence

Student Preface

The International Space University (ISU) was founded in April 1987 as a non¬
profit, non-governmental institution. It was created with the objective of
becoming the world's leading centre for educating and training tomorrow's
space professionals. The ISU Summer Session Program brings together
international space experts from academia, industry, and government to
educate students in multidisciplinary and advanced issues in space
development in a ten week format. The design projects carried out by the
students during the session have two purposes: first, to provide learning in
international teamwork on problems requiring a multidisciplinary and
multicultural approach, and second, to yield published results that can be
influential in the world-wide space community.

This year's summer session was held in Vienna, Austria, and this report
outlines the effort of one of its two groups of students. The team, composed
of 53 professionals from 18 countries, brought to the project a variety of
experiences, educations, and interests, from the societal through to the
scientific, from the theoretical through to the applied. The members of our
group used varied styles of problem solving, ranging from the ambitious and
unconstrained to the more limited and immediately achievable.

Our mandate was to use an international perspective to examine present and
planned activities in solar-terrestrial science and applications, critically
review current goals, investigate new organisational schemes, develop
innovative mission concepts and define a comprehensive baseline project that
represented a realistic alternative or follow-on to the projects now being
considered in space agencies.

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Faculty Preface

At each ISU summer session the students carry out one or more design
projects. Their purpose is to give experience in intercultural and
multidisciplinary teamwork and at the same time to generate results that can
be influential in the world beyond ISU and useful to the students in their
later careers. At ISU 96 the two projects were about remote medical activities
and solar-terrestrial science and applications, named by the students DOCC
and Ra respectively. Of the 104 members in the ISU class of 1996, fifty-three
people from eighteen countries and all ISU academic disciplines chose to
work on Ra. This document delivers their results.

The charge to the student team was for them to use an international
perspective to examine present and planned activities in solar-terrestrial
science and applications, critically review current goals, investigate new
organisational schemes, develop innovative mission concepts and define a
comprehensive baseline project representing a realistic alternative or follow-
on to the projects now being considered in space agencies.

Recognising that the realm of Sun-Earth interactions is huge and diverse, the
students had to make choices using their own judgement as to what they
could achieve in a short project. They developed a Strategic Framework
containing near, mid, and far term activities for both science and applications
and analysed those that they believed most promising. They used
information and advice from their faculty and teaching assistants plus that
contributed by other members of the ISU community and visiting experts.
They made effective use of the new information facility provided bv the
World Wide Web. y

The students' decisions on what to analyse and report, what to treat by
reference, and what to omit from the project were entirely their own. We, the
faculty and teaching assistants for this project, are honoured and proud to
have been associated with this energetic, disciplined and creative group of
students and we commend their results to the reader.

James D. Burke

California Institute of Technology
Jet Propulsion Laboratory

Chantal Lamontagne
Carleton University

George Scoon
ESTEC

European Space Agency

Gregory Mallory
University of New Brunswick

Executive ^“flf
Summary |||

In this report, we set out a framework for pursuing solar science and appli¬
cations. As a guiding charter, we have chosen the following mission state¬
ment:

Through an international perspective, we will
explore and document strategies which will increase
our understanding of the Sun and its effects, and
help us apply solar knowledge for the benefit of
humankind.

Ra Team Mission Statement

The timing is fortuitous.

The ESA Science Programme
Committee (SPC) will be meeting in
November 1996. After this meeting,
the Call for Ideas for the M4 mission
(part of the Horizon 2000 Plus pro¬
gramme) will most likely be released.
The M4 has presently been reserved
for a mission concentrating on the
Solar System.

Also in the immediate future, the
Inter-Agency Consultative Group for
Space Science (IACG) will likely
begin the process of choosing its next
focal project. Currently, they have

been co-ordinating the International
Solar Terrestrial Physics Program
(ISTP).

Furthermore, NASA is planning to
bring its Sun-Earth Connections
Roadmap to the American space sci¬
ence community for assessment. That
meeting is set for the summer of 1997
at Woods Hole, Massachusetts.

We encourage the wider community
to investigate the contents of our full
report. Much of it has taken the form
of recommendations for the future,
and many ideas await your discovery
within.

Report Overview Diagram: The hieroglyphs were found using the URL of Laurent Wacrenier,
Nam en hieroglyphes, http://yoko.ens-cachen.fr:8080/hiero, accessed August, 1996.

vi

• Ra- Thf* ^un Fnr ^ripnrp ariH Humanitv

- The International Situation

The global political environment global space infrastructure, used by

within which space activities take both developed and developing

place has been changed by a variety nations, points to an immediate need

of economic, social, and technological for improved solar warning and fore¬
factors. This altered paradigm has casting capabilities. The political

created both obstacles and opportuni- environment recognises these eco-

ties for solar exploration and applica- nomic needs, resulting in an

tions. enhanced opportunity for develop¬

ments in solar warning and forecast-
The end of the Cold War has had the ing.
most far-reaching implications for

national space activities. Deep and There has been an international trend

integrated co-operation in space toward greying the line between the

between the United States and Russia basic and applied sciences. This grey-

is no longer a political taboo, opening ing has the potential to enhance the

up a whole new array of international cohesion of the scientific community

co-operative opportunities. by diminishing traditional rivalries

Conversely, the loss of competitive between speciality disciplines. The

Cold War rationales has been a pri- convergence is also notable for the

mary driver of the decreasing nation- movement toward interdisciplinary

al space budgets in both the United science missions, and the current cli-

States and Russia. These same mate is favourable toward joint sci-

decreasing budgets stimulate < ence and applications endeavours,

increased national inter-agency

operation and co-ordination. 'ThiSr;,^ , ? The future of solar exploration and
trend toward greater collaboration|'||5> ^ applications will be determined large-
presents an opportunity for a mult i-k ly by how well the relatively low
lateral co-operative effort in solar^®*^ budgetary priority of solar and helios-
exploration and applications. pheric physics and solar warning and

forecasting services is overcome. The
The respective technological levels of combination of diminishing national

spacefaring nations are no longer dis- space budgets, increased opportuni-
parate. Although economic competi- ties for co-operation, and growing
tion between spacefaring nations has technological capabilities has led to a
partly supplanted the old political sustainable emphasis on smaller,

competition of the Cold War, less modular, networked spacecraft with

commercial sectors, such as space sci- prioritised objectives. Disciplinary
ence, have experienced enhanced co- cohesion, inter-agency co-ordination.

operation because of mutual payback
opportunities and decreased concern' I
about disproportionate or unilateral *§
technology transfer

The economic risks
global knowledge
ous solar phenom
new heights Th
amount and lev

international co-operation, applica¬
tions rationales, and smallsat technol¬
ogy offer a combination of effective
to sustain and even increase
exploration and applications

;; j i. y* ffi

yze & . HwSiSs 5 :3S‘ !

Executive Summary • vii

A Strategic

One of our goals in the report is to
develop a Strategic Framework for
solar science and applications, and
from that programmes for the Near-
Term, Mid-Term, and Far-Term. This
Strategic Framework provides an
integrated approach to solar explo¬
ration and application, as illustrated
in the figure below. Three time
frames are defined as follows:

• Near-Term: Focuses on programmes
that are achievable within the next
few years (1996 to 2000). Elements
tap into current capabilities and
programmes; they also seek to
improve management and co-opera¬
tive structures in preparation for the
future.

• Mid-Term: Focuses on more ambi¬
tious programmes, some requiring
technology development, with
implementation times in the first
decade of the next century (2001 to
2010).

• Far-Term: Focuses on the period
from approximately 2011 to 2020
(and beyond) and is characterised
by higher-risk, advanced techno¬
logy, and/or integrated pro¬
grammes.

The elements of the Near-Term pro¬
gramme are primarily political and
managerial in their scope, in keeping
with the Near-Term philosophy of
building on existing capabilities.
Central to this programme is the cre¬
ation of a "Working Group for
International Solar Exploration and
Applications" (WG ISEA). We envi¬
sion the WG ISEA as a forum for co¬
ordinating and planning the many
solar missions that individual nations
have proposed for the next decade,
while preserving their independent
sources of support. These missions
tend now to be rather random. Other

WorkltigGroup for International Solar Exploration^ Application

Increase awareness

Coordinate data

Improve forecasting models

| SAUNA: Heliocentric science platform {Solar threat monitoring & early warning system^ * 6 *

:T6lfdW

■xux..

Ouster

recovery

World forecasting system

Global climate effects
monitoring > ;

Heliospheric observer

0 .

Stereo coronal imaging

follow-on programme
follow-on programme

jCo-wlinated FIRE]

SuiadeprabfSc:

Integrated science 4c applications program

viii • Ra: The Sun For Science and Humanity

Framework

parts of the programme may not be as
ambitious but can have profound
implications. The sharing of science
data, for example, may produce
synergistic results and lead to better
solar environment forecasting models.
Overall, the Near-Term programme
lays a foundation for the projects of
the later parts of the Strategic
Framework.

We propose several mission flight
opportunities in the Mid-Term period.

A stereoscopic solar imaging system
is envisioned to fulfil the high priority
science objective of understanding the
corona, as is a heliocentric near-Sun
science platform (which we have fli
named SAUNA). The corona is cur- ; V h
rently scheduled to be probed by the
combined Russian-US FIRE mission.

These missions will be supported by a
new global heliospheric observation
system (possibly one of the stereo
observation platforms), since SOHO
may have expired and not been
replaced by the time it is needed to
support FIRE and other missions. We
envision a continuously operating
solar threat monitoring and early
warning system, perhaps one involv¬
ing near-Sun platforms that build on
the technology demonstrated by
SAUNA. This system will mark the
beginnings of a solar applications
system, an idea central to Ra. Finally,
we envision that humanity will be
taking serious steps toward the estab¬
lishment of human lunar outposts or

future. Building on the foundations
created earlier — better forecasting
models, data co-ordination, increased
solar awareness, and the WG ISEA
(whose international activities will
have continued and expanded in
importance) — we envision an inte¬
grated programme for space science
and applications. This integrated pro¬
gramme may have combined plat¬
forms, or it may share common
resources (such as spacecraft bus
designs, or a communication system
to relay data from a new generation
of solar spacecraft). The space threat
monitoring and early warning system
begun earlier should be mature
|err0ugh by this time to create a global
forecasting system, one that provides
§|!plS||f|? *° developing nations,
•"‘‘•'irmore, applications will begin
is on solar benefits, such as the
beginnings of space solar power
plants. Finally, as we look back on
years of integrated data, we see these
data, combined with new long-term
missions, enabling scientists to study
the relationship between the Sun and
the Earth's climate.

We believe the Ra Strategic
Framework is significant because it:

• is a coherent plan over time.

• relies on existing and planned pro¬
grammes, and benefits from them.

--

Mars exploration; in which case,
study of solar radiation's effects on J|fe
tissue will be essential t^lhe design
of these missions. In the

• considers the political and econom-
^ environment, including future
nds, and seeks to shape that en-
^n^wment for the advancement of
:ience and applications.

ijllk.

science and applica-

To guide the development of the Ra
Strategic Framework, we defined sci¬
entific and applications objectives. For
our primary areas of scientific inter¬
est, we chose the corona, the solar
wind, the Sun's effect on the Earth,
and solar theory and model develop¬
ment. For secondary areas of scienti¬
fic interest, we selected sunspots, the
solar constant, the Sun's gravitational
field, helioseismology and the galactic
cosmic rays. We stress the importance
of stereoscopic imaging, observations
at high spatial, spectral, and temporal
resolutions, as well as of long dura¬
tion measurements. Further explo¬
ration of the Sun's polar regions is
also important, as shown already by
the Ulysses mission.

From an applications perspective, we
adopted three broad objectives that
would derive complementary inputs
for the Strategic Framework. These
were to identify and investigate: pos¬
sible application spin-offs from sci¬
ence missions, possible solar-terrestri¬
al missions dedicated to a particular
application, and possible future appli¬
cations that require technology devel¬
opment. The Sun can be viewed as
both a source of resources and of
threats. Our principal applications
focus was that of threat mitigation,
by examining ways to improve solar
threat monitoring and early warning
systems.

We compared these objectives to the
mission objectives of past, current.

and planned international solar mis¬
sions. Past missions (1962-1980) seem
to have been focused on improve¬
ment of scientific knowledge, using
multiple instrument spacecraft. A ten
year gap followed this period, during
which the results from previous mis¬
sions were analysed and solar study
programmes were prepared in inter¬
national organisations. Current mis¬
sions (1990-1996) focus on particular
topics such as the corona, solar flares,
and coronal mass ejections. In
planned missions, Sun/Earth interac¬
tions and environmental effects of
solar activity are becoming more
important. The corona is the centre of
interest of almost all planned mis¬
sions. It seems that no international
long-term strategy has yet been
adopted. For these plans the number
of necessary future missions can be
reduced and the onboard instrumen
tation can be optimised by perform¬
ing a comparative analysis.

The study of the corona must be done
from different observing locations,
orbits closer to the Sim, and by differ¬
ent means. The Cluster mission
replacement is in progress; however,
if the replacement is not implemen¬
ted, the ISTP programme will fade
after 1998. Furthermore, the physics
of the Sun's interior should be
emphasised more in the Mid- and
Far-Term programmes. Finally, more
emphasis should be placed on moni¬
toring space weather and forecasting
Sun/Earth interactions.

The continued expansion of solar
understanding will necessitate
research rationales that include both
basic and applied scientific objectives.
To properly integrate these rationales,
a single forum for solar exploration
and applications co-ordination and
planning is optimal. The Ra Strategic
Framework calls this forum the
Working Group on International Solar
Exploration and Applications (WG
ISEA). To take full advantage of cur¬
rent events in space science, the WG
ISEA should be formed before the
Summer 1997 NASA Woods Hole
Sun-Earth Connections Roadmap
meeting.

The programmatic means by which
the WG ISEA achieves its internation¬
al collaborative objectives should be
flexible to maximise the political sus¬
tainability of the effort. The WG ISEA
should include a Mission Co-ordina¬
tion Group to synthesise co-ordina¬
tion and data sharing between nation¬
al solar science and applications mis¬
sions outside, with, and beyond the
International Solar Terrestrial Physics
programme (ISTP). To supplement
the inevitable gaps in solar observing
capabilities that will still exist, the
WG ISEA should also form a Mission
Planning Group to recommend a
strategic framework for solar explo¬
ration and applications that takes
advantage of existing, cheap plat¬
forms, such as university mini-satel¬
lites, for quick response solar observa- „

1 _J__■ l 1 1

is also a key to reducing the cost of
solar system exploration. To take
advantage of this economic opportu¬
nity while realising its political reali¬
ties, the WG ISEA should include an
engineering group for the internation¬
al design of reference models for solar
spacecraft. This Reference Model
Design Group provides a first step
towards realising the benefits of inter-
nationakco-operation in space explo-
ration beyond the co-ordination of
scientific data acquisition and data

Increa se d understanding of solar and
helid^sigrib physics will generate
advances in solar forecasting models,
and current national plans to consoli¬
date agency-level solar warning and
forecasting resources will incorporate
these advances. Existing international
solar warning and forecast data distri¬
bution networks like the International
Space Environment Service will feed
data into these forecasts, but the
advances needed to make solar warn¬
ings and forecasts relevant to poten¬
tial users will require capital invest¬
ment in hardware, especially in
instruments placed between the Earth
and Sun. National solar warning and
forecasting plans should look abroad
for opportunities to co-ordinate the
deployment of dedicated but nation¬
ally discrete solar warning spacecraft.
Meeting user needs will provide hori¬
zontally integrated commercial
opportunities within the larger gov-

i i- - luiuuw vvxLiu.li me larger gov-

tion or solar instrument technology '^eminent space warning and forecast
demonstration. ^^p^yices. A solar warning spacecraft

y~.. . ,, • •* V* JMlBMso la <ely be the first operational

Discrete national hard|fe||gnbibuf ; g ; :t^^S^^ffih a ^ -

tions to internationa’
the political enviro
activities. The use
common spacecr

_ , jrace endeavour outside

„ and fe%»ology demon-

r.a»’ ’v.'-, r yVl^'’i a *>. *

^ an impor-
anity's

ill Whets and Funding

* 5 ■ f There is a market transformation tak¬

ing place from the public sector to a
combination of the public and private
sectors. Our vision is to support this
transformation and to expand and
fully use existing and potential mar¬
kets. Our research has found three
major markets for Ra:

• Entertainment and education mar¬
kets can be served by converting
the Ra scientific results. This will
increase the public awareness
about the Sun and its effect on the
Earth and human life.

We expect these markets to evolve as
shown in the figure below.

• Space environment forecasting is
an increasing market, and the next
ten years will see it increase from
$100 to $200 million U.S. annually.
Potential markets are influenced by
insurance companies and financial
institutions. These markets are
sensitive to failures of telecommu¬
nication satellites and energy sup¬
pliers.

Increasing public interest in the Ra
programme will likely increase the
availability of governmental funding.
We recommend further studies.

Space agencies are interested in solar
science and space environment fore¬
casting. Improved measurements and
models of the space environment will
benefit both manned and unmanned

. The science market will expand as space programmes mid thereby con-
Ra increases the benefits through stitute a ground for funding.

augmenting scientific and techno¬
logical knowledge. This increase
will help develop and implement
solar illumination and solar heating
infrastructure systems. Including
these in buildings and transporta¬
tion systems has the potential to
significantly influence the well¬
being of the global population.

There is a trend toward joint ventures
between universities and industry.
The universities' research is relevant
to industry, and industry funds part
of it. We see a trend where Sun activ¬
ities are moving from being research
driven to product/service driven.

A Near-Term

past. We call for the Working Group
for International Solar Exploration &
Application (WG ISEA) to be started
in the Near-Term. To help advance
the Mid- and FarrSerm programmes
through to fruifionyfwe advocate
increasing awareness of solar science
and solar-terrestrial connections,
thereby fostt||ng*support beyond the
scientific community. Finally, in the

Programme

Description

Cluster recovery

A replacement for the Cluster programme and direct new Cluster mission toward Ra's
objective

Improve forecasting
models

Perform correlation studies; innovative acquisition of new forecasting models

Co-ordinate science and
other data

Continue ground-based observations; create an international data centre; research with and
co-ordinate science data; co-ordinate future planning of independent groups

Working Group for
International Solar
Exploration and

Application (WG ISEA)

Incorporates science and applications interests from government and private sectors; submits
to government agencies speriBc recommendations for actions necessary for the fulfilment of
the solar exploration and application strategic plan, while encouraging independent
complementary efforts

Increase awareness of solar
science and Sun-Earth
interaction

Develop a "common language" for solar science and applications; work with planetariums
and museums; educators via WWW; correlation study on satellite anomalies, ground power
station anomalies and solar activity

Actively incorporate
existing technology
initiatives

Examples include: Japan Nereus, ESA TRP (esp. Theme 10) and GSTP, NASA New

Millennium, University Small Sat, Clementine, DC-XA, Commercial bus

Each part of the Near-Term pro¬
gramme is relatively low in cost and
either builds upon existing systems
and infrastructure or incorporates
modest developments. We believe
that the recommendations are realistic
and play an important role in realis¬
ing important science and applica¬
tions objectives. They also provide a
foundation for the projects described

in the Mid- and Far-Term pro¬
grammes.

Near-Term programme, we support
actively incorporating existing tech¬
nology initiatives.

To build on existing solar observation
instruments (namely SOHO) and to
continue with a logical sequence of
solar observation satellites, we recom¬
mend recovery of the Cluster pro¬
gramme. As we believe space envi¬
ronmental forecasting will become
more important to the space commu¬
nity in the Mid- and Far-Term, we
recommend immediate work on
improving forecasting
amount of archived
grow and additional
satellites are launch^
ordination of and
both the new dati

The most significant suggestions are
two correlation studies: one to estab¬
lish the relationship between solar
activity and satellite anomalies, and a
second to evaluate the accuracy of
current solar activity forecasting mod¬
els. These are interrelated and each
serves, in the Near-Term, to get the
ications objectives "off the

.uestoj

;a^^^^^m£^comp^n^its of the Near-

rl**' * ' T * 1

A Mid-Term Programme

The Ra Mid-Term framework aims to:

• provide a solar science programme
to address fundamental issues of
solar physics.

• improve the capability for solar
applications, and do so in co-ordi¬
nation with the science pro¬
gramme.

The second objective is served by a
transient phenomena monitoring and
early warning system, and a small but
important human dosimetry payload.
The latter is clearly needed for the
safety of manned interplanetary mis¬
sions, and as such must fly before a
crewed expedition to Mars or a lunar
base become reality. The stereoscopic
mission will open the third dimension
for solar physics, flying moderately
capable remote sensing instruments at
1 AU on small spacecraft buses, shar¬
ing heritage with existing small satel¬
lites. This will also serve as a precur¬
sor to an operational stereoscopic
solar event prediction and early
warning system. The SAUNA mis¬
sion aims to send a medium-sized sci¬
ence payload to a moderately close
heliocentric orbit inside that of
Mercury, at about 0.2 AU. This mis¬
sion will provide long-term high reso¬
lution monitoring of the solar disk in

the extreme ultraviolet and of the
corona in white light. Stereoscopy
and contextual measurements will be
possible when the data are combined
with those from observations made
on or near the Earth. SAUNA will
also act as a technology demonstrator
for subsequent long-term missions in
closer orbits such as a heliosynchro-
nous/polar constellation system.
SOHO is showing the value of long¬
term heliospheric measurements from
an orbit not significantly nearer the
Sun than the Earth. Although it will
probably remain operational until
2004, the planning of a replacement
must start now if new and outstand¬
ing questions about the Sun are to be
investigated effectively. The new
platform should aim to reduce mis¬
sion cost while improving capability,
since SOHO itself is clearly a "mon¬
ster mission" using large-scale 1980's
technology. The currently proposed
joint Russian-US FIRE mission, a
simultaneous dual-spacecraft close
flyby of tiie Sun to investigate the
corona, is included in Ra's Strategic
Framework. The dual mission is of
far higher scientific value than if only
a single spacecraft were flown.

The major components of the Mid-
Term programme are summarised in
the following table:

Programme

Description

SAUNA: a heliocentric,
near-Sun science platform

Ion-propelled single spacecraft to 0.2 AU heliocentric orbit. 5 yr. mission duration

Solar threat monitoring and
early warning system

Heliocentric orbiters; Other options included: L4/L5 tripwire and solar wind event imaging
and tracking

Stereoscopic corona
imaging system

Small remote sensing platforms at LI, L4 and L5

New heliospheric
observing platform

Extended SOHO mission, then smaller follow-on

Co-ordinated FIRE

Mission: Russian Plamya
and U.S. Solar Probe

Dual spacecraft dose flyby mission to 4 Rg and 10 Rg

Human radiation studies
on host spacecraft

Tissue-equivalent dosimeter measuring direct radiation and secondary radiation from
shielding

xiv • Ra: Thp Sun For Science and Humanity

A Far-Term

The Far-Term programme of the Ra
Strategic Framework is designed to
build upon the experience gathered
during the Mid-Term programme. We
assume that more ambitious and
higher-cost projects are possible in the
Far-Term, providing that these are
balanced by a proportionally
increased economic viability in terms
of commercial exploitation and direct
benefits to society.

Propulsion: further impn
in ion engine performam
opment of prototype s<
vehicles for the inner so.
further research into a<
cepts like mass drivers

Power: high efficiency heat resis¬
tant solar arrays

Programme

Description

Integrated solar science and
applications programme

pWTOof coSnS apphcab0nS; a PP Ucation P rot °tyP* sensors on science

Small suicide probes

Wide range of concepts available

World-wide space
environment forecasting
system

Charactensbcs include: distributed provides information to developing nations, integrates
military, evil, commercial data; independently maintained in partidpating natoT^

Preliminary space solar
power applications

Prototype space-based solar power station for small-scale distributed use

Monitoring the Sun's effect
on Earth's climate

Long-term space-based observation programme to monitor solar output and Earth's climate

Integrated solar science and applica¬
tions programmes would succeed in
reducing cost through co-operation in
areas of common interest and through
exploiting available opportunities.
Small suicide probes would explore
the acceleration and heating in the
solar corona by means of in situ mea¬
surements. A world-wide space envi¬
ronment forecasting system would
offer benefits to all humankind.
Preliminary solar power applications
would be instrumental in exploring
ways to solve the imminent global
energy crisis on Earth. Monitoring the,
solar constant and its effect on the j
Earth's climate would allow study of
tne impact of vanatior^fj^the solar
output on the Earth'

In order to succeed
the following tec
ments will be

• Materials: high-temperature cera¬
mics and alloys

• Electronics: radiation hardened
high-temperature electronics, more
powerful small lasers

• Communications: optical commu¬
nication techniques

• Guidance, Navigation and Control:
autonomous interplanetary navi¬
gation techniques (e.g. based on
planetary ephemerides), increased
on-board intelligence

■ p Launchers: low-cost access to orbit
, ffl^ means reusable launch

^SwfMiateles

rised in

The Ra report is a call to action.
Knowledge of the Sun is vital to us as
humans and to our planet. Our star
deserves our attention and study.

The global political environment
within which space activities take
place is changing for a variety of eco¬
nomic, social, and technological rea¬
sons. The current international situa¬
tion presents both obstacles and
opportunities for solar exploration
and applications. This situation is
ideal for the introduction of Ra.

We present in our report a Strategic
Framework for pursuing solar science
and applications. From this
Framework a programme emerges for
the Near-Term, Mid-Term, and Far-
Term. We believe the Ra Strategic
Framework is significant because it:

• offers coherency over time.

• utilises, benefits from, and adds to
current programmes.

• harmonises with our political and
economic environment.

• integrates solar science and appli¬
cations.

• capitalises on global talents and
resources.

By defining and analysing objectives,
we give impetus and focus to the
Strategic Framework. We have identi¬
fied potential markets and sources of
funding.

We recommend that a Working Group
for International Solar Exploration
and Applications (WG ISEA) be estab¬
lished immediately. The WG ISEA
would:

• ensure that a Strategic Framework
is put into action.

• synchronise independent efforts in
different countries.

• facilitate the interaction between
science and applications.

• help to combine the output into
products useful on a global scale.

The time is opportune, ideal for the
introduction of our ideas into the
space science and applications com¬
munity. Having in place a Strategic
Framework dedicated to solar science
and applications, and forming a small
but broadly-based international WG
ISEA would prove most beneficial.

We hope that our report will help to
make this happen.

Table of Contents

Introduction.

1 • 1 Mission Statement. . .

1.2 Strategic Framework. 2

1.3 Report Organisation. . .

1 4 Organisational Diagram. . .

The Ra Strategic Framework. 4

2.1 Overview of Programme Elements. . .

2.2 Factors Considered in Developing the Strategic Framework.... . g

2.3 Implications. .

Political & Economic Environment. ^ 4

3 2 * e S ' a ® e for totemali onal Co-operation: Criteria and Modelling I’’ ” 16
(WG ISEa” " g . r0 “ P .° n . International Solar Exploration and Applications

.. 07

3.3 Data Dissemination Principles for the Ra Project .

3.4 Organising for Solar Warning and Forecasting ZZZZZZZZ .38

3.6 Solar Research and Forecasting in the Context of Ru^ian Space Poli™. t

3.8 Co“;r 1“ and ^ ^ ^ S,rategiC F “ - = ^

Our View of the Sun. . .

4.1 Studying our Sun. . .

4.2 The Sun as a Star. . .

4.3 Interplanetary Space. . 64

4.4 The Sun-Earth Interactions. . 76

4.5 Effects of the Sun on Earth, Humans and Technology. . II

4.6 The Sun as a Resource. .

Objectives & Requirements. . .

5.1 Science Objectives and Priorities in the Ra Strategic Framework. 99

5.2 Applications Objectives and Priorities in the Ra Strategic Framework III: ,01

5.4 Scenarios. ..

5.4.2 Spacecraft Fleet and Trajectory.Z.Z.. .

5.5 Recommendations on Requirements . .

Technology Challenges and Issues. ZZZ. .

6.1 Solar Environment. ..

6.2 Payload Instrumentation. . .

6.3 Orbit and Trajectory Definition.’ ’ .. 117

6.4 Propulsion. . .

6.5 Power Systems. . .

.. 140

6.6 Structures and Materials.

6.7 Thermal Control Technology Challenges

6.8 Guidance, Navigation and Control.

6.9 Communications.

6.10 Command anc

6.11 Opportunities
Standardisation in

Market and Funding Issues.

7.1 Markets for ..

7.2 Project Funding.

7.2.2 Private Funding.

7.2.3 Combination of Private and Governmental Funding.

7.3 Marketing.

Near-Term Programme.

8.1 Overview.

8.2 Replace Cluster.

8.3 Improve Forecasting Models.

8.4 Co-ordinate and Apply Science Data..."

8.5 The Near-Term Role of the Working Group for International Solar

Exploration & Application (WG ISEA).

8.6 Increasing Awareness...

8.7 Actively Incorporate Existing Technology Initiatives.

8.8 Conclusions.

Mid-Term Programme.

9.1 The SAUNA Mission.

9.2 Solar Threat Monitoring and Early Warning Systems.

9.3 Solar Stereo Mission.

9.4 New Heliospheric Observing Platform.

9.5 The Fire Mission...

9.6 Mission to Determine Biological Radiation Effects.

9.7 Mid-Term Costing.

I Data Handling...

for Spacecraft Commonality, Modularity and
Future Solar Science and Applications Missions

Far-Term Programme.

10.1 Integrated Solar Science and Applications Programme

10. 2 The Suicide Probe...

10.3 World-Wide Space Environment Forecasting System..

10.4 Preliminary Solar Power Applications.•.

10.5 Monitoring the Solar Effect on the Earth Climate.

10.6 Costing of the Far-Term Programme.

10.7 Conclusions.

Conclusion.

Appendix A - Overview of Sun Related Missions.

Appendix B - SAUNA Mission Data.

Appendix C - Technology Challenges and Issues.

Appendix D - Costing.

Appendix E - Existing and Proposed Early Warning Systems

References.

...145
... 153
...156
...163
.... 169

....177
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....187
.... 191

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.195

.195

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.209

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.248

.. 251

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.259

.260

.. 262

.266

.267

.268

.270

.270

..275
..303
..309
.. 321
_ 333
„. 335

List of Figures

Figure 1.1 Report Overview Diagram. 4

Figure 2.1 Strategic Framework Overview. 6

Figure 2.2 Strategic Framework development process. 6

Figure 2.3 Strategic Framework Inter-relations Matrix. 14

Figure 3.1 Organigram of the Working Group on International Solar Exploration and

Applications. 30

Figure 3.2 The "Triple I" Model.

Figure 4 .1 The interior of the Sun. ., c

Figure 4.2 The H-R diagram. 66

Figure 4.3 The path and position of our Sun. 66

Figure 4.4 Principal zones in the solar. 67

Figure 4.5 Modes of oscillation in the Sun.. ^

Figure 4.6 The variation of temperature with height in the solar atmosphere.70

Figure 4.7 Coronal mass ejections. 73

Figure 4.8 Composite eruption model. 7 g

Figure 4.9 Spiral IMF lines frozen into a radial solar wind expansion at an average speed of
4UU Km/S. wjrj

Figure 4.10 Interaction with bodies in the Solar System. 7 g

Figure 4.11 A schematic showing a magnetic cloud modelled as a toroidal magnetic flux

r °P e ..... 79

Figure 4.12 Three-dimensional cutaway view of the magnetosphere showing currents fields
and plasma regions. ' g^

Figure 4.13 The Earth radiation belts. 02

Figure 4.14 Typical ionospheric electron density profiles. g 4

Figure 4.15 The Sun viewed as a resource. 96

Figure 4.16 Space tourism, the next step. 97

Figure 5.2 Space Region Classification. ^ 30

Figure 6.1 Jupiter Swing-by with 90° Inclination Change. 130

Figure 6.2 Resonant Venus Flyby. ^

Figure 6.3 Venus-Mercury Flyby followed by Electrical Propulsion. 132

Figure 6.4 Top View of Venus - Mercury Flyby. 132

Figure 6.5 Venus - Mercury Flyby Av Requirement. 132

Figure 6.6 Top View of Direct Injection. 133

Figure 6.7 Av Requirement for Direct Injection. 133

Figure 6.8 Top View of Direct Injection and Electric PropulsionCombination. 133

Figure 6.9 Av Requirement for Direct Injection and Electric PropulsionCombination......... 133

Figure 6.10 The Five Lagrangian Points. 134

Figure 6.11 Polar Eccentric Orbit 0.2-1.4 AU using Solar Sailing. ZZ.'.'.'Z *135

Figure 6.12 Schematic Representation of a Solar Sail with a Tilting Angle for Decreasing the
Velocity of the Spacecraft. ° j 39

Figure 6.13 Closed Steam Cycle. 1^2

Figure 6.14 Stirling Engine Principle. 147

Figure 6.15 P-V- and T-S Diagram of the Ideal Stirling Cycle.143

Figure 6.16 Peltier Element.

Figure 6.17 Electrodynamic Tether.

Figure 6.18 Specific Power .

Figure 6.19 Specific Costs..

Figure 6.20 Frame and Unified Volume Spacecraft Structure. 147

Figure 6.21: Specific Properties of Typical Aerospace Materials...149

Figure 6.22 Specific Strength Versus Temperature for Metal- and Ceramic-Matrix ^

Composites.".

Figure 6.23 Temperature of a typical heat shield and the heat flux as a function of the

distance from the Sun (steady state).- I 55

Figure 6.24 Advantages and Disadvantages of Ka Band Over X Band.166

Figure 6.25 Typical Block Diagram of a Deep Space X Band Transponder.166

Figure 6.26 Factors Relating to Spacecraft Autonomy. 176

Figure 7.1 End-to-end chain of users of space environment prediction.188

Figure 7.2 Estimation of market evolution over time. 191

Figure 7.3 Market demand as a function of marketing effort. 49

Figure 8.1 A low-cost alternative for Cluster recovery?.I 97

Figure 9.1 Overview of SAUNA Mission Objectives.210

Figure 9.2: Functional Flow Block Diagram for the SAUNA Mission.211

Figure 9.3: Low Thrust Trajectory from 1.0 AU to 0.2 AU Circular Orbit:.212

Figure 9.4 SAUNA Spacecraft - Selected Configuration.214

Figure 9.5 Spacecraft Orientation. 215

Figure 9.6 The Thrust Axis Can Be Pitched To Provide Out-of-Plane Av.215

Figure 9.7 Propulsion System Layout. 217

Figure 9.8 Thermal Model of Heat Shield, MLI, and Instruments.. 220

Figure 9.9 The instrument temperature as a function of the distance from the Sun for various

dissipation levels.

Figure 9.10 The Ion Thruster Radiator Temperature as a Function of the Distance from the

Figure 9.11 SAUNA Programme Timeline.228

Figure 9.12 Logical Sequence for the Study.229

Figure 9.13 Connection between solar phenomena and effects on the ground and on space ^

systems.

Figure 9.14 Orbital configuration of option A. zo °

Figure 9.15 Orbital configuration of option ..233

Figure 9.16 Orbital configuration of option ..234

Figure 9.17 Optimisation of heliocentric distance.237

Figure 9.18 Solar event advance warning time.238

Figure 9.19 Communications link geometry.240

Figure 9.20 Mission cost breakdown.242

Figure 9.21 Cost Breakdown structure of Space segment.243

Figure 9.22 Cost breakdown structure.253

Figure 9.23 The cost as a function of payload mass and the distance from the Sun.254

Figure 9.24 Cost breakdown for SAUNA mission.255

Figure 9.26 Cost breakdown for Ra Application mission.- 256

Figure 10.1 Conceptual Future Network of Sun-Orbiting Spacecraft.260

Figure 10.2 The Temperature of the Heat Shield Near the Sun.263

Figure 10.3 The Required Av as a Function of Perihelion.- 264

YY •

Ra: The Sun for Science and Humanity

Figure 10.4 The quotient Mass (propellant)/Mass (Total Initial Spacecraft) as a Function of
the Perihelion. 264

Figure 10.5 Solar Heat and Light Distribution System for Buildings. 268

List of Figures • xxi

List of Tables

Table 2.1 Near-Term Programme (1996 to 2000).

Table 2.2 Mid-Term Programme (2001 to 2010).

Table 2.3 Far-Term Programme (2011 to 2020 and Beyond).

Table 2.4 Primary Science Objectives.

Table 5.1 Past Missions: General Objectives.

Table 5.2 Past Missions: Measurements.

Table 5.3 Current Missions: General Objectives.

Table 5.4 Current Missions: Measurements.

Table 5.5 Planned Missions: General Objectives.

Table 5.6 Planned Missions: Measurements.

Table 5.7 Needs and Measurements.

Table 5.8 Spacecraft Fleet and Trajectory.

Table 5.9 Environment and Subsystems.

Table 6.1 In Situ Measurement Types.

Table 6.2 Remote Sensing Measurements.

Table 6.3 Characteristics of the Sensors.

Table 6.4 The Model Payload for a Future Mission.

Table 6.5 Characteristics of the JPL EUV/VIS Remote Sensing Instrument.

...7
... 8
...9
..12

104

105

105

106
106
107

109

110
111
118

119

120
. 121

123

Table 6.6 Summary of Physical Characteristics of Remote Sensing Instruments.124

Table 6.7 Velocity Increments for Various Ecliptic Heliocentric Orbits.129

Table 6.8 Typical Performance Data for Various Chemical Propulsion.136

Table 6.9 Advantages and Disadvantages of Various Chemical Propulsion Technologies. 137

Table 6.10 Solar Array Parameters. 141

Table 6.11 RTG Advantages and Disadvantages. 141

Table 6.12 Power Source Comparison.- 144

Table 6.13 Primary Battery. 145

Table 6.14 Secondary Battery.- 145

Table 6.15 General Requirements.- 146

Table 6.16 Frame Spacecraft Structure. 147

Table 6.17 Unified Volume Space. 147

Table 6.18 Material Characteristics of Light Metal Alloys.150

Table 6.19 Accuracy, FOV, and Sensitivity (mv) of Some Star Sensors.161

Table 6.20 Performances of Some Reaction Wheels.162

Table 6.21 Performances of Some Reaction Wheels (Cont.).162

Table 6.22 Summary of Space Radiation Environment and Their Effects on CMOS Electronic
Devices. 172

Table 9.1 Evaluating 3-Axis vs. Spin Stabilisation.

Table 9.2 UK-25E thruster.

Table 9.3 Solar Arrays.

Table 9.4 Thermal Requirements for the Components

. 213
..217
218
- 219

Table 9.5 Overview Mass and Power of SAUNA Thermal Control Subsystem
Table 9.6 SAUNA Mission Budget.

Table 9.7 Early Warning System- Customer Requirements.

Table 9.8 Early Warning System qualitative trade-off matrix.

Table 9.9 Payload Estimates for Heliocircular Array Spacecraft.

Table 9.10 Av values for several heliocentric distances.

Table 9.11 Spacecraft mass distribution.

Table 9.12 Total mass launched vs. distance from the Sim.

Table 9.13 Comparison of propellant masses to propulsion systems considered for 1.0

Table 9.14 Ra Applications cost matrix.

Table 9.25 Cost comparison of SAUNA with similar missions.

Table 9.27 Cost comparison of Ra application mission with similar missions.II

.222

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.229

.235

.236

.237

.239

.239

AU to
.241

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.257

Acronyms

"Triple I"

H-sun

ACE

AC RIM

ACS

ADCS

AOCS

APS

AU

CCD

CDF

CDM

CDR

CEOS

CIS

CME

CMOS

CNES

Co-I

COSPAR

CR

CSA

CSDS

CSEF

CSIRO

CSW

DARA

DC

DIPS

DLR

DOC

DOD

DOE

DOI

DSN

EDAC

ELDO

EOM

ESA

ESOC

ESRIN

Interagency/International Interface organizational model for solar
warning and forecasting services
Sun gravitational constant.

Advanced Composition Explorer

Active Cavity Radiometer Irradiance Monitor

Attitude Control System.

Attitude Determination and Control System.

Attitude and Orbit Control System
Active Pixel System.

Astronomical Unit (1 AU=1.5 10 8 km) Mean distance between the

Sun and the Earth

Coupled Charge Device

Common Data Format

Code Division Multiplexing

Critical Design Review

Committee for Earth Observing Satellites

Commonwealth of Independent States

Coronal Mass Ejection

Complementary Metal-Oxide-Semiconductor.

National Center for Space Studies (France)

Co-Investigator
Committee on Space Research
Commissioning Review
Canadian Space Agency
Cluster Science Data System

Committee for Space Environment Forecasting (U.S.)
Commonwealth Scientific and Industrial Research Organisation
Committee for Space Weather (U.S.)

German Space Agency
Direct Current

Dangerous Interplanetary Plasma Structures

Deutsche Forschungsanstalt fur Luft- und Raumfahrt e. V.

U.S. Department of Commerce

U.S. Department of Defense

U.S. Department of Energy

U.S. Department of the Interior

Deep Space Network

Error Detection And Correction

European Launcher Development Organisation

End Of Mission (Review)

European Space Agency
European Space Operations Center
European Space Research Institute

EUMETSAT

FAC

FAST

FDIR

FDM

FGF

FY

GCC

GCR

GEO

GGCM

GGS

GIC

GMS

GNC

GPS

GSOC

GSTP

HESI

HGA

HRG

i

IACG

IC

IFOG

IKI

IMAGE

IMEWG

IMF

IMP

IMU

INPE

IP

IPS

IRU

ISAS

ISEE

ISES

ISES

ISL

ISPM

ISRO

ISS

ISTP

IZMEM

IZMIRAN

JSOC
LI, L4, L5
L4,L5
LAN

European Meteorological Satellite Organisation

Field Aligned Current

Fast Auroral Snapshot Explorer

Failure Detection, Isolation and Recovery

Frequency Division Multiplexing

Fluctuating Geomagnetic Field

Fiscal Year

Gore-Chemomyrdin Conference
Galactic Cosmic Rays
Geostationary Earth Orbit
Geospace General Circulation Model
Global Geospace Science
Geomagnetically Induced Current
Geostationary Meteorological Satellite (Japan)

Guidance, Navigation and Control.

Global Positioning System
German Space Operating Centre
General Support Technology Programme
Fligh Energy Solar Imager
High Gain Antenna
Hemispherical Resonator Gyro
Inclination of the orbital plane.

Inter-Agency Consultative Group
Integrated Circuit
Interferometric Fibre-Optic Gyro.

Soviet Space Research Institute (now with Russia)

Imager for Magnetopause to Auroral Global Exploration
International Mars Exploration Working Group
Interplanetary Magnetic Field
Interplanetary Monitoring Platform
Inertial Measurement Unit.

National Institute for Space Research (Brazil)

Inter Planetary

Interplanetary Plasma Structures
Inertial Reference Unit.

Institute for Space and Astronautical Sciences (Japan)

International Sun Earth Explorer

International Solar Energy Society

International Space Environment Service

Inter Satellite Link

International Solar Polar Mission

India Space Research Organization

International Space Station

International Solar-Terrestrial Physics Programme

IZMIRAN Electro-Dynamic Model

Institute of Terrestrial Magnetism, Ionosphere, and Radio Wave
Propagation

Joint Science Operations Center
Lagrange Point 1/4/5, Libration Point 1/4/5
Stable Lagrangian points.

Local Area Network

LEO

LGA

MAMS

MEO

METEOSAT

MMH

MMS

MOA

MOU

MSM

MSM

MSTI

mv

NASA

NASDA

NMP

NOAA

NSF

NSSDC

NSWC

NSWP

NTO

OFCM

OPAL

OSO

OSS

OSTP

OTH

PDR

PE

PI

PRR

QMPP

R&D

RAM

RCS

Re

RF

Rs

RSA

RTG

RW

RWC

SAC B

SAMPEX

SAPPHIRE

SAUNA

SCF

Low Earth Orbit
Low Gain Antenna

Modular and Multifunctional and Systems (one of NMP's
Integrated Product Teams)

Medium Earth Orbit.

European Meteorological Satellite Programme
Monomethyl Hydrazine

NASA Goddard Space Flight Center's Multimission
Modular Spacecraft
Memorandum of Agreement
Memorandum of Understanding
Magnetospheric Specification Model
Minimum Solar Mission.

U.S. Air Force Miniature Sensor Technology Integration
satellites and program

Visual magnitude of the stars. It defines star sensor sensitivity.
National Aeronautics and Space Administration
National Space Development Agency of Japan
Jet Propulsion Laboratory's New Millennium Program
National Oceanic and Atmospheric Administration (U.S.)
National Science Foundation (U.S.)

National Space Science Data Center
National Space Weather Council (U.S.)

National Space Weather Program (U.S.)

Nitrogen Tetroxide

Office of the Federal Coordinator for Meteorology (NSF)
Stanford University's Orbiting Picosat Automatic
Launcher

Orbiting Solar Observatory

Office of Space Science (part of NASA)

Office of Science and Technology Policy (U.S.)
Over-The-Horizon radar
Preliminary Design Review
Pluto Express.

Principal Investigator

Preliminary Requirements Review

Quantitative Magnetospheric Predictions Program

Research and Development

Random Access Memory

Reaction Control System

Mean Earth Radius = 6371 km

Radio Frequency

Solar Radius

Russian Space Agency

Radioisotope Thermal Generator

Reaction Wheel.

Regional Warning Centers

Satelite de Aplicaciones Cientiffcas

Solar Anomalous Magnetospheric Particle

Stanford Audio Phonic Photographic Infrared Experiment

Solar Adjacency Using a New Approach

Smaller, Cheaper and Faster

vvvi • Ra: The Sun for Science and Humanity

SCM

Standardisation, Commonality and Modularity

SEC

Space Environment Center (NOAA, formerly SEL)

SEC

Sun-Earth Connection

SEE

Single-Event Effects

SEL

NOAA Space Environment Laboratory (NOAA now SEC)

SEP

Sun-Earth-Probe

SESC

Space Environment Services Center (U.S.)

SEU

Single Event Upset

SiC

Silicon Carbide

SMEI

Solar Mass Ejection Imager (Phillips Lab)

SOHO

Solar Heliospheric Observatory

SOI

Silicon On Insulator

SOLRAD

Solar Radiation

SOS

Silicon On Sapphire

SPC

Science Programme Committee (part of ESA) IACG Inter Agency
Consultative Group

SPF

Sun Protection Factor

SPP

SAUNA Predevelopment Programme

SQUIRT

Stanford University's Satellite QUIck Research Testbed

SRR

System Requirements Review

SSC

Sudden Storm Commencements

SSR

Solid State Recorder

STA

Science and Technology Agency of Japan

STSC

Star Tracker Stellar Compass.

STSP

Solar-Terrestrial Science Programme

SWRI

Southwest Research Institute

TAOS

Technology for Autonomous Operational Survivability

TDM

Time Division Multiplexing

TEPC

Tissue Equivalent Particle Chamber

TID

Total Ionising Dose

TIMED

Thermosphere, Ionosphere, Mesosphere, Energetics and Dynamics

TRACE

Transition Region And Coronal Explorer

TRP

Technology Research Programme

TTC

Tracking, Telemetry and Command

U.S.

United States

UK

United Kingdom

USAF

United States Air Force

uv

Ultraviolet

WAN

Wide Area Network

WG

Working Group

WG ISEA

Working Group on International Solar Exploration and Applications

WWA

World Warning Agency (ISES)

WWW

World Wide Web

XTE

X-Ray Timing Explorer

Chapter 1

Introduction

Through an international perspective, we will explore and
document strategies which will increase our
understanding of the Sun and its effects, and help us
apply solar knowledge for the benefit of humankind.

''One of the least understood objects in the solar system is our star the S,m " tk

Keeping in mind the mission statement, we produced a pltn of acttai^our-Ser'
ramework, an outline providing possible direcHon for the future of solar exploration. 8 *

a B s e ked e indude n8 ' hiS “ Wi * ‘° pause and “P°" *e past. Questions often

What reasons do we have for pursuing these investigations?"
How did we arrive at our present situation?"

"What have we learned thus far?"

"Where have we failed before, and why?"

"What is our current situation?"

"What would we like to do next?"

"What are we able to do next?"

, inrtiiziHnak and as society, we humans have entered into and

Culture and daily life are shaped by the Sun, perhaps more than by any other na ra
body.

stars by seeking to learn more about our own.

The inhabitants of any neighbourhood areMertwined gm in

donated by y its n stan e Our Sun S its own unique way of communicating. It attracts. It
emits. It broadcasts. It expels.

_ , j c tl p nlanet on which we live takes all of this solar input and

neighbourhood.

We ourselves as members o< ~r —^ arealso

KSSr* community. There is also much to be lost whenever we
delay the next step in our involvement.

mmimm

Lvandng o?h“g S our continued involvement, the latter being an essent.al
ingredient which may influence the quality of life on our planet. Earth.

1.1 Mission Statement

- .„ in the Droiect the students of the Solar Probe Design Project Team agreed to adopt
Early in the project, ,, . . d Dro bl e m like this, it is useful to provide some

smte'menf Memory 8 focuses the objectives, and indicates the mam

priorities.

From our mission statement we derived the following goals and objectives:

• To explore and document the science and applications needs for the future;

To develop a Strategic Framework for solar science and applications and
rom at a program for near-term, mid-term, and far-term missions; and

‘ in po,icy - fundi "S'

1.2 Strategic Framework

From the mission statement, we focused our attention towards defining a strateev for

Our approach for developing this framework consisted of various steps First we
researched possible objectives that would satisfy current scientific and application needs

Ae past ThT;: d ; heSe ? ie ^ ves t0 "l OSe of various solar ™si°ns thJt have flown in
the past, that are currently flying or that are planned for the future The resulting

objectives formed the basis for new solar missions, which were then evaluated in liaht if

•he poimeal, budgetary, and technological challenges .ha, they may“ he S r

that' t Jt 7 he r ! su t is a “Aprehensive program fo/science and applkaHons

exploration missions. “ ““ ^ a ^ut

1.3 Report Organisation

Our report reflects the importance of the Strategic Framework we present The eleven

nolSr? m h K 1S reP ° rt C °, U d be grOU P ed int0 three P ar ‘s that supportthis framework the
that IS’ andeconomical environment that sets the stage for the framework; the issues
^ ^ description of the ™

Political and Economic Environment

In this sect,on, consisting mainly of chapter 3, we present a broad picture of the political
and economical environment that affect most of the decisions currently made about Sar
exploration pro,ects. We provide the reader with concrete examples Ulusffa^no1
concepts presented and examine lessons learned from these case studies We explore the
“Ifr* international eo-operafion through solar science and appXahons
udy models for this co-operation and propose a new organisation, the Working Group

for International Space Exploration and Applications (WG ISEA) responsible for
overseeing this co-operation and establishing a data dissemination structure.

Issues that Shape the Strategic F ramework

The chapters in this section [chapters 4, 5, 6, and 7] provide a background, rationale and
issues that mould the Strategic Framework. First, we present a description of so ar
science as we view it today. The Sun is presented not only in a scientific context but also
in a historic and societal context that should provide a general view of the Sun to the
reader After the foundation has been set, we describe the scientific and applications
objectives that drive the need for solar exploration. In this description we discuss pas
Sd present needs from a broad perspective. Applications objectives are p»»ted «*
only in light of the threats posed by the Sun, but also m the opportunities that the Sun
may present for potential technological advances. Once these science and applicatio
objectives are identified, we present the technological and economic issues that constrain
these objectives and that influenced our decisions on shaping the framework. We discuss
challenges, and how these challenges may be overcome.

Strategic Framework Missions

Chapters 8, 9, and 10 of the report discuss the missions that make up the structure of our
Strategic Framework. These missions are grouped based on chronological distribution
near term missions, mid-term missions, and far-term missions. They are categorised
based on the use of existing technology and capability, as well as on their availability.
Assembled together, these missions constitute a complete plan for solar exploration that
spans several decades of scientific investigation and opportunities for applications.

1.4 Organisational Diagram

The interaction between the three sections described above is represented in figure El
and at the beginning of each subsequent chapter. The figure provides an overview of the
entire report and helps to place each chapter within the context of the Stra eg
Framework. As we go through each chapter, that chapter will be highlighted m grey.

l Our View of the Sunyj

L Needs & Objectives^

arkeung^^undin^

Q Jwm*1 \

c > a <=» I

Strategic Framework,

Near-Term

Mid-Term

1 Political & Economic*
V Environment J

Far-Term

tut

. Conclusions

Fie 11 Report Overview Diagram: The Heiroglyphes were found using the URL

8 ‘ ' of Laurent Wacrenier "Norn en hiQro g 1 y p h e s ,

http://yoko.ens-cachan.fr8080/hiero, accessed August 1996

Chapter 2

The Ra Strategic
Framework

s mentioned in the Introduction, one of our goals in this report is to "develop a Strategic

. ^f?^ wor ° r s 2 , ar sc ' ence an< ^ applications, and from that a programme for Near-Term
Mid-Term Far-Term Missions". This Strategic Framework provides an integrated

“ 3nd appliC3ti0n ' aS illuStrated %^e 2.1. Thref time

1. Near-Term: Focuses on programmes that are achievable within the next few
years (1996 to 2000). Elements tap into current capabilities and
programmes; they also seek to improve management and co-operative
structures in preparation for the future.

2. Mid-Term: Focuses on more ambitious programmes, some requiring
technology development, with implementation times in the first decade of
the next century (2001 to 2010).

3. Far-Term: Focuses on the period from approximately 2011 to 2020 and
beyond, and is characterised by high-risk, advanced technology, and/or
integrated programmes.

In this chapter, we present the Ra Strategic Framework: its programme elements, the logic
behind its development, and special implications. We developed the Strategic
Framework by consulting science and application experts; developing and assessing
objectives, examining instruments and technical capability; considering policy and
business concerns; and conceiving and assessing scenarios. Our approach is illustrated in
igure 2.2. A similar analysis is being conducted by NASA's Office of Space Science: the
Connection (SEC) Roadmap [Sun-Earth Connection Roadmap, WWW] Unlike
the SEC Roadmap, the Ra Strategic Framework is international and less concerned with
recommending specific programmes than with focusing the direction of exploration and

applications (the former is beyond the scope of the report). Also, we avoided
investigating the Earth's magnetosphere—this area is too complex for an adequate
investigation given our schedule.

WoAtog Croup far Sol^ E»rlotrti°n * __|

SAUNAi HcUomttrlc pl.rfon. | ScUr «<»«»*« * «dT .y*- p"*"*” (

Fig. 2.2

Strategic Framework development process.

2.1 Overview of Programme Elements

In this section
Framework. A
8 , 9, and 10.

we present and discuss the programme elements of the Ra Strategic
more detailed description of the individual elements is found in chapters

Table 2.1 Near-Term Programme (1996 to 2000).

Programme

Cluster

recovery

Objectives
Complement to

SOHO and ground-
based observations

Description

A replacement for the
Cluster programme and
direct new Cluster mission
toward Ra's objective

Rationale

Utilise all of the existing
work done for the original
Cluster toward what Ra
team believes to be the most
pressing concerns

Improve

forecasting

models

Improve space

environment

forecasting

Perform correlation studies;
innovative acquisition of
new forecasting models

Current operational
forecasting models are old
and empirical; better
models will save
degradation and
replacement cost

Co-ordinate
science and
other data

Make use of all past
and current data

Continue ground-based
observations; create an
international data centre;
research with and co¬
ordinate science data; co¬
ordinate future planning of
independent groups

Other research communities
may be interested in solar
data, easier data access
provides more time for
actual research

Working

Group for

International

Solar

Exploration

ana

Application
(WG ISEA)

An international
forum for the
planning, co¬
ordination, and
implementation of an
international effort in
solar exploration and
applications

Incorporates science and
applications interests from
government and private
sectors; submits to
government agencies
specific recommendations
for actions necessary for the
fulfilment of the solar
exploration and application
strategic plan, while
encouraging independent
complementary efforts

Changing global paradigm
for space science and
applications points to the
advisability of combining
resources across both
national boundaries and
science vs. applications
disciplines. We believe WG
ISEA is the most efficient
and expedient
organisational forum to
enable this merger

Increase
awareness of
solar science
and Sun-Earth
interaction

Increase awareness
among: general
public, space
community, power
companies

Develop a "common
language" for solar science
and applications; work with
planetariums and
museums; educators via
WWW; correlation study on
satellite anomalies, ground
power station anomalies
and solar activity

Maintaining funding will
require a basic public
understanding; science, as a
"public good , should be
snared; establishing a
correlation between space
weather and satellite
anomalies will motivate
further investigation/
interest

Actively

incorporate

existing

technology

initiatives

Continue with
efficient technology
development

Examples include: Japan
Nereus, ESA TRP (esp.

Theme 10) and GSTP,

NASA New Millennium,
University Small Sat,
Clementine, DC-XA,
Commercial bus

Matches post Cold War era
frends; logical progression
into the future

2.1.1 Near-Term Programme

The elements of the Near-Term Programme, presented in table 2.1, are primarily political
and managerial in scope, in keeping with the near-term philosophy of building on
existing capabilities. Central to this programme is the creation of a "Working Group for
International Solar Exploration and Application" (WG ISEA). We envision the WG ISEA

nJl° r r CO '° T ? r at T S and Planning the man y Solar missions ^at individual

nations have proposed for the next decade while preserving their independent sources of

support. As discussed in chapter 5, these missions currently tend to be rather random

Other parts of the programme may not be as ambitious but can have profound

The Ra Strategic FrampwnrV

• 7

implications: the sharing of science data, for example, may pr^uce synergistic results
and lead to better solar environment forecasting models. Overall, the Near ' T ^
Programme lays a foundation for the projects of the later parts of the Strateg c

Framework.

Table 2.2. Mid-Term Programme (2001 to 2010).

Programme
SAUNA: a
heliocentric,
near-Sun
science
platform

Obiectives

High resolution
coronal and surface
imaging; in situ solar
wind measurements;
technology
demonstrator

Description

Ion-propelled single
spacecraft to 0.2 AU
heliocentric orbit. 5yr.
mission duration

Rationale

Affordable ($200M) science
mission and demonstrator
of survivability near Sun;
precursor to heliocentric
constellations

Solar threat
monitoring
and early
warning
system

Measure position,
velocity of southward
interplanetary
magnetic fields

Heliocentric orbiters;

Other options included:
L4/L5 tripwire and solar
wind event imaging and
tracking

Initial dedicated space
environment system;
selected option most
compliant with identified
potential customers

Stereoscopic

corona

imaging

cvQtem

Magneto-
hydrodynamics of
corona

Small remote sensing
platforms at LI, L4 and L5

First stereoscopic mission-
low cost but high return-
opening the third
dimension

j y oici11

New

heliospheric

observing

nlatform

Helioseismology,
solar atmospheric and
coronal stuaies, solar
wind monitoring

Extended SOHO mission,
then smaller follow-on

Maintenance of long-term
observation and monitoring

yj A Cl VI KJl l a i

Co-ordinated
FIRE Mission:
Russian

Plamya & U.S.
Solar Probe

Heating of the corona
and acceleration of
solar wind

Dual spacecraft close flyby
mission to 4 R$ and 10 R$

Low-cost close flyby
mission with finely targeted
objectives

Human
radiation
studies on host
spacecraft

Determine radiation
risks for humans in
interplanetary space
and requirements for
protection

Tissue-equivalent dosimeter
measuring direct radiation
and secondary radiation
from shielding

Essential precursor for
human Mars exploration or
lunar base; could be a
show-stopper

2.1.2 Mid-Term Programme

The elements of the Mid-Term Programme are presented in table 2.2. We propose several
missions in this time period. A stereoscopic solar imaging system is envisioned to fulfil
the high priority science objective of understanding the corona, as is a heliocentric near-
Sun science platform (which we have named "SAUNA"). The corona will also be probed
by a combined Russian-U.S. FIRE mission. These missions will be supported by a new
global heliospheric observation system (possibly one of the stereo observation platforrm)
since SOHO may have expired by the time it is needed to support FIRE and other
missions [Randolph, 1996]. More significantly, we envision a continuously-operating
solar threat monitoring and early warning system, possibly invoking near-Sun platforms
that build on the technology demonstrated by SAUNA. This system will mark the
beginnings of a solar application system, an idea central to Ra. Finally, we hope that
humanity will be taking serious steps to the establishment of human lunar outposts or
Mars exploration; in which case, study of solar radiation's effects cm humans will be
essential to the design of these missions. In summary, the Mid-Term Programme
elements represent a maturing of solar science and the beginnings of solar applications.

a « i? a . Thp Sun for Science and Humanity

Table 2.3 Far-Term Programme (2011 to 2020 and Beyond).

Programme

Objectives

Description

Rationale

Integrated
solar science
and

applications

programme

Reduce costs by co¬
operation in areas of
common interest and
by exploiting free
opportunities

Options: science
“piggybacking" on
applications; application
prototype sensors on
science platforms; use of
common buses

Solar science and
applications have common
elements; an integrated
programme spreads risk
and provides synergistic
benefits

Small suicide
probes

Explore acceleration
ana heating of corona
by direct sensing

Wide range of concepts
available

Understanding of coronal
phvsics is of high scientific
value

World-wide

space

environment

forecasting

system

Enhance the benefits
of space environment
forecasting for
humankind

Characteristics include:
distributed, provides
information to developing
nations, integrates military,
civil, commercial data;
independently maintained
in participating nations

Political, social and
commercial interests
ultimately converge in the
maximum availability of
early warning systems

Preliminary
space solar
power
applications

Explore ways to solve
the imminent global
energy crisis

Prototype space-based solar
power station for small-
scale distributed use

Solar power represents a
"next generation"
application

Monitoring
the Sun's
effect on

Earth's climate

Understand the
impact of variations
in the solar output on
the Earth's climate

Long-term space-based
observation programme to
monitor solar output and
Earth's climate

Co-ordinated programme
allows long-term data to be
gathered so that potential
correlations can be
uncovered

2.1.3 Far-Term Programme

The elements of the Far-Term Programme look toward the more distant future. Building
on the foundations created earlier — better forecasting models, data co-ordination
increased solar awareness, and the WG ISEA (whose international activities have
continued and expanded in importance) — we envision an integrated programme for
space science and applications. This programme may have combined platforms, or it mav
share common resources (such as spacecraft bus designs or a communication system to
relay data from a new generation of solar probes). Also, the space threat monitoring and
early warning system begun earlier should be mature enough by this time to create a
global forecasting system, one that also provides benefit to developing nations.
Furthermore, applications will begin to focus on solar benefits: the beginnings of space
solar power plants. Finally, as we look back on years of integrated data, we see these data

(combined with new long-term missions) enabling scientists to study the Sun's influence
on harth s climate.

2.2

Factors Considered in Developing the Strategic Framework

We considered several factors while formulating the Strategic Framework. Among those
highlighted below are: policy drivers (political and economic); science objectives'
applications objectives; past, current, and planned missions; technology; programme
element inter-relationships; orbital vs. flyby missions; and our vision for the future.

2.2.1 Policy Drivers

The Strategic Framework is shaped by political and economic factors that transcend the
scientific objectives, applications needs, and technological opportunities for solar
observation. In this section we delineate these factors, including an overview of the
!mpact these "policy drivers" have had on the Strategic Framework. Further information
on the politico-economic environment can be found in chapter 3.

2.2.1.1 Post Cold War Environment

The principal policy driver for the Strategic Framework is the evolving Post Cold War
environment for space activities. This environment possesses inherent benefits and
drawbacks. For example, it provides opportunities for scientific co-ordination between
former adversaries on solar missions like FIRE (see Mid-Term Strategic Framework)
while depriving space activities of their former national security rationales and funding
levels which limits Strategic Framework recommendations in the near-term. Many of
the policy drivers listed below will refer to the Post Cold War environment as their
definitive paradigm.

2.2.1.2 Convergence of International Technology Levels

Less than two decades ago, the technological capabilities of the Soviet Union and the
United States easily outstripped those of the other spacefaring nations. Today, the gap
between the technology pools of the former superpowers and those of the other
spacefaring nations has drastically narrowed. Although this shortening gap fosters
national and commercial competition in space technology development, it also promotes
success when international co-operation in solar observation missions is undertaken. On
a level technological playing field, partners are able to offer more resources and benefits
to each other, and the costs of international co-operation are reduced through the
common technical literacy of the partners. International co-operation is also no longer
primarily limited to scientific data co-ordination. Converging international technology
levels make co-operation in spacecraft and mission engineering more likely, and the
Strategic Framework takes advantage of this by emphasising the need to include
engineers in an international solar working group. The Strategic Framework also takes
advantage of converging technology by setting an objective for the engineers in t is
international solar working group: the production of common, spacecraft system designs
to serve as world-wide baseline reference models to make solar observation missions

more affordable.

2.2.1.3 Global Nature of Solar Threats

Dangerous solar phenomena and their interaction with the near Earth space environment
and the Earth's upper atmosphere and magnetic fields transcend national boundaries.
Although the damage to specific human resources may be nationally local, rarely is the
damage from a solar incident limited to one nation's resources. The rising world-wide
technology pool (described in section 2.2.1.2) and the increasing number of spacefaring
nations (described in section 2.2.1.5) put more and larger human resources at the mercy of
solar phenomena. Understanding these phenomena requires data from nations around
the world Though international scientific and solar forecasting organisations do exist to
ensure that this data is exchanged and disseminated, the improvement of current solar
forecasting models and solar warning systems would benefit more from international co¬
operation and co-ordination at the level of space hardware. The Strategic Framework
favours organisational and technical solutions to space warning and forecasting that go
beyond mere data sharing.

2.2.1.4 Flat or Declining Space Agency Budgets Among Developed Countries

Without Cold War rationales for space activities, space agencies throughout the
developed world have found their budgets levelling out or declining with time. Solar
research already a low priority in many space agencies, will suffer if actions are not taken
to counteract its budget priority and its available resources. Declining space agency
funds require missions that fit within small budgets, require various solar science
disciplines to prioritise their objectives with one voice, and require solar observers and
forecasters to multiply their resources by going outside their agencies and nations. The

Strategic Framework highlights spacecraft with low budget ceilings, a means for
organising solar science and applications disciplines internationally and solar data
acquisition and modelling resources outside naHonal space agencies „ the academic
commercial and military spheres. '

2.2.1.5

Emerging Space Capabilities in Developing Countries

The developing world is becoming more reliant on space activity to create the
infrastructure needed for prosperity and to monitor the externalities
continued' Z ) econo ™ c S r °wth. While these fragile capabilities are essential for

techn^raf ft Iff °^T ' deV f°^. countries ma y lack th ? resources, both material and
frnm A ' * effectlvel Y P rotect their nascent space and terrestrial technology systems

wa^in anSe dT S ° ar P henomena ’ ^tegration of the developing world's needs in solar

StrategkF a ramewor C k § ° rganiSati ° nS neCessitates international co-operation in the

2.2.1.6

Increasing Co-operation Between National Agencies

Co-operation in solar observation and forecasting among national space agencies
weather agencies, science and technology development agencies, and militaries is
required by the flat or declining budgets each is being subjected to in the Post Cold War
environment. Previous budgets allowed these national actors to duplicate early solar
observation and forecasting capabilities. New budgets drive them to co-operate to
preserve old capabilities and necessitate co-operation to create new ones. The Strategic

a l a Zn°J P n ntS ° Ut # °PPf rtun ; ties to shar * data, human resources, hardware, and costs
at the national level to further solar science, warning and forecasting.

2.2.1.7 Interdependence of Solar Science and Space Warning and Forecasting

Applications 6

Expanding basic knowledge about the Sun and its interaction with the Earth's
magnetosphere and atmosphere will be crucial to refining solar forecasting models

sterns 11 Th Sola ^ r f earch wil1 also find applications in sola? warning

systems^ The S rategic Framework has attempted to expand, rather than narrow, the
links between solar science and solar warning and forecasting.

2.2.1.8 Trend Towards Interdisciplinary Science Missions

Because solar science is a low budget priority for most space agencies, the Strategic
Framework has sought out opportunities for solar observation wherever they may be
found. These opportunities include missions that piggyback solar sensors on other
spacecraft and missions that use hardware developed for other uses to perform solar
observation for science or forecasting.

2.2.1.9 Emergence of Smallsat Technology

Smaller, faster, cheaper concepts have driven missions in the Strategic Framework to
consider current smallsats for new solar observation missions in the near- and mid-term
and to design high technology, low mass, standardised smallsats for mid- and far-term

missions. Constellations and commonality are two important concepts that drove Ra
mission selection. r d

2.2.2 Science Objectives and Priorities

The Strategic Framework concentrates on the high priority science objectives identified in
chapter 5 [section 5.1] and summarised in table 2.4 below. These objectives concentrate

on the corona, solar wind, and the Sun's influence on Earth's climate. Accomplishment of
these objectives requires long duration observations from appropriate vantage points.
Hence, our emphasis on stereoscopic observations, global solar observation, heliocentric
orbital platforms, and occasional solar probes. These efforts must be co-ordinated to
achieve maximum benefit: co-ordination of missions and of the resulting data.

Table 2.4 Primary Science Objectives.

Primary Objective

Investigation Areas

To understand the physical
processes leading the Sun to
emit plasma structures and high
energy particles that are
potential threats to humans and
technology.

• Heating mechanism of the corona

• Formation of coronal holes

• Emergence of the slow solar wind

• Relationship of fast solar wind to coronal holes

• Causes of and underlying physical principles of solar flares

• Causes of the acceleration of particles to very high energies

• Release of coronal mass ejections (CME's)

• ProDaeation of CME's in the interplanetary medium

To understand the physical
processes which may lead the

Sun to influence our climate.

• Cause of changes to the solar constant

• Long-term variations in the solar constant

2.2.3 Applications Objectives and Priorities

Given the applications objectives discussed in chapter 5 [section 5.2], we focused our
priorities on one application: solar threat monitoring and early warning. Such an
application is in its infancy, and a mature market does not exist. Therefore, the creation
of a viable market is a primary concern in developing the Strategic Framework. Since
funding is also limited, existing resources must be maximised: as sources of data and as a
means to improve forecasting models. By laying a solid foundation in the Near-Term
Programme and by taking realistic steps in establishing initial capabilities, we believe a
viable, self-sustaining system will follow.

2.2.4 Past, Current, and Planned Missions

In developing the Strategic Framework, we also examined past, current and planned
solar missions [see chapter 5, section 5.3]. Several conclusions resulted from this
comparative analysis:

. There is no global co-ordinated plan for solar exploration, although there is some
activity, such as the International Solar-Terrestrial Program (ISTP);

• A solar applications programme is lacking; and

. Study of the corona is a hot topic: it was studied by eleven out of twenty past
and current missions (since 1962), and seven out of the eleven planned missions
plan to collect more data. The high priority given to coronal study as a science
objective means that continued observation from different spatial, spectral,
and/or temporal perspectives is necessary.

2.2.5 Role of Technology

Solar missions benefit from advanced technology in three ways:

L " 3n be madC S , ma ! ler and m ° re effective ' hereby reducing costs

( bmallsats were previously discussed in Section 2.2.1.9); and

2. Innovative thermal protection technology can help protect close-to-the-Sun

missions (e.g., 0.4 AU) which face a harsh environment (e.g., temperatures
communication interference); and 8 P '

3. Mission hardware requires high Av's to get into their proper orbits.

Thus, the use of advanced, "leading edge" technology is advocated in the Strategic

Near-Term' telnol f * *5f deSig " S ° f individual Programme elements. In the
m?PW ! T 0m effortS such as ESA ' s Technology Research Programme

NASA rjpw vrif r • eme 10 j T d General Support Technology Programme (GSTP),
NASA s New Millennium, and the U.S. Clementine programme should be exploited

Te?m C p reqUinn8 V f u advanced technologies, however, should be placed in the Far-
m Programme of the Strategic Framework, allowing time for these technologies to

tTm?' [Worden ^ W * Sh ° Uld 6XpeCt 0nl y " one ^acle at a

nnMW ^ 1996]. Otherwise, delays and cost overruns will result, endangering not

only that particular project but possibly other elements of the Strategic Framework. §

2.2.6 Programme Element Inter-relationships

The Strategic Framework is programme in time. Not only did we divide it into three
consecutive periods: near, mid, and far-term. We also desired that individual programme
elements fol owed a logical progression (see figure 2.1 at the beginning of this chapter)

below 11 TV 6 f ^ etW n en pr0gramme elements are further illustrated in figure 2.3
• is figure also illustrates that some programmes are complementary: FIRE for

exampie requires a heliospheric observer (like SOHO) for instrument calibration and a
global solar reference [Randolph, 1996]. a a

2.2.7 Orbital vs. Flyby Missions

Achieving our science objectives requires long-term observation. Hence, the Strategic

Framework favours heliocentric orbital missions over short duration flybys. However

sometimes critical data cannot be gained without directly sensing the phenomenon of

in erest. Therefore the Strategic Framework still needs to consider proSeHn the

Strategic Framework, the 'suicide probes" following FIRE are placed in the Far-Term

Programme - after we received the results from FIRE and heliocentric missions when

technology may better support near-Sun probes (e.g., thermal protect and

communications improvements), and when a science/application heliocentric system

m T Up r , thes * probes (eg ' actin S as a communications relay or as a "piggyback
mother ship" to reduce costs). v my

2.2.8 A Vision for the Future

The Ra Strategic Framework is a focused path to the future. In developing that path we
asked ourselves where we wanted it to lead. Common responses included "integrated "
global and the next step." The Far-Term Programme allowed us to formalife these
ideas which ranged from an integrated science and application programme and a
programme that benefits all regions and aspects of the globe (e.g., developing countries
an understanding global climatic change) to the beginnings of using the Sun as a
resource (e.g., space solar power stations). Some of these elements are not very visionary;

2.3 Implications

We believe the Ra Strategic Framework is significant because it:

• Is a coherent plan over time;

• Relies on existing and planned programmes and benefits from them,

• Considers the political and economic environment, including future trends,
and seeks to shape that environment for the benefit of solar science and
application;

• Integrates solar science and applications, showing how one benefits the
other;

• Is an international framework that capitalises on global talents and
resources; and

• Seeks to provide global benefits.

Additional study is required for specific programmatic decisions. We hope, however,
that the Ra Strategic Framework will have a positive influence on increasing our
understanding of the Sun and its effects, helping to apply that knowledge for the benefit

of humanity.

tdiaj Chapter 3

Political & Economic
Environment

lu^1^2i eC inai C J^ omxn> is a p T erful force shapin * ,he nalure and

^ Ch p 2/ we overview ed some primary policy considerations that
e the general direction and configuration of the Ra Strategic Framework This

deempd ™ 1SSUeS more de P th and add considerations of policy topics

deemed important for the success of the Ra Strategic Framework. We begin the chafer

by setting the stage for international cooperation with an analysis of paft and existing
examples of cooperation in space. We continue to build on that foundation with a
recommendation for the structure of an international cooperative forum for solar and
heliospheric science and applications, the Working Group on Internarional Solar

of fntema^onaf da^dis^ issues conce ™"8 the successful implementation

S^™ramework. 0 " StmCtUreS "* ^ aSSeSSed 38 ** relate * a

Through cooperative associations like the Working Group, knowledge about solar
processes and their influence on the space environment is increased and this knowledge
has practical applications m space environment forecasts, forecasts that can hefp
governments and industries mitigate or even prevent damage to terrestrial and space
resources from dangerous solar phenomena. Although an array of agency national and
international resources exists to aid the pursuit of /viable solar wamtag anS forecast
semce, space weather forecasting exists at an infantile state of development, requiring
the measured marshalling of information and hardware resources to improve the

YodelsY S ° lar “ ar '’ in 8 s a "d ^que current and future organizational

models for consol,dating and increasing the capabilities of solar warning and forecYt

ervice resources. Drawing recommendations from this analysis, we then present a new
rgamzational synthesis, the Inter-agency/International Interface ("Triple I") Model for
future solar warning and forecasting organizations. Chapter 3 concludes with a
consideration of Russian contributions and participation in international solar forecasting
organisations and a review of international and national contracting arrangements

The ideas and issues addressed in chapter 3 create a textual structme ( * at * e

reader to more readily appreciate the environment that shaped the Ra Stra g
Framework and this report s remaining chapters.

3.1 Setting the Stage for International Co-operation: Criteria and
Modelling

In order to establish an effective international framework for solar and heliospheric
science and applications, it is important to first define the environment within which the
framework must function, and then describe some means by wh,ch **
maximise its chances of survival and success in such an environment. With these
considerations in mind, the Ra team has evaluated six examples of Internationa
operation in space activities, and drawn upon these examples for lessons we can learn
and apply to our formulations for Ra. The following cntena and project analyses, then,
provide us with a foundation upon which the Ra team can build an international co¬
operative framework for solar exploration and applications.

3.1.1 Criteria for Solar Science Co-operative Frameworks

International co-operation in space has taken on many forms since the Soviet Union first
launched Sputni/on October 4, 1957. It is difficult to speak of success in many cases,
however without first defining what success means. Success for a scientist is th * ret “™
of useful data- for an engineer it is a fully operational spacecraft; and for politicians
t “ess is often defined less tangibly in terms such as technology transfer, political
influence, and economic return. All definitions of success are both valid and vital for
their mutual achievement is essential to maintain overarching support for a projec or
programme. However, the varied faces of success are often prob ematic because in many
cases conflict can occur if the goals of the partners (at all levels are not at least
compatible if not complementary. The attainment of the overall success of an
international project can very often be judged as a product (at least partially) of the
political and managerial frameworks under which the endeavour functioned. The
purpose of this section is to define both constraints which must be met if an international
co-ordinating framework for Ra is to have some chance at fulfilling its mission, and some
"optimisation means" that may give the framework a better chance of doing so.

3.1.1.1 Constraints

Political and managerial frameworks inherently function within a certain set of
constraints Most obviously, these constraints are restrictions im P 0S ^f ^ the overa
leeal and political structure of the involved nations. For instance, when NASA engages m
!f co-operative venture, the Memorandum of Understanding signed between the parties
tifThere is any) always includes a clause similar to, "subject to the availability of funds .
This is due to the political structure of the United States, which precludes NASA from
obligating Congress to appropriate funds. However, structural constraints such as this
are unavoidable, and the space activities of nations alone are not likely to precipitate
fundamental alterations of the national political frameworks involved^ The polihcal and
Wal structures under which space activities take place transcend beyond individual
sectors (such as space), and thus respond rather inflexibly to the needs of national space

activities alone.

Taking this into consideration, it is helpful to have some working definitions of the most
pressing constraints within which any international organise!,onal framework for Ra

must function.

distinct i^^ompatib^^urther^eseTa^ona^f^mus^bTbo^c^arly^suled^in^da^
simple terms and well-communicated ,„ the nationa. conunules tavXd ^

mmunities to effectively influence the national funding mechanisms in their favour

SMEEJffir

most often. Upon entering into a co-operative framework, minimum sources of funding

mH35

process. 7 PPr ° Ve P r0 S rammes ln order to minimise their vulnerability in the funding
gc onomic Return - No matter how attractive Ras scientific and application potential or

muh^^^

zsr x.s^s&r rantee each partidpa "* *

ompames. Financially speaking, this makes a match between a country's contribution to
ESA and the contracts it gets back, and is therefore referred to as >ste retou » pXv

ban fers C ^ ra T W 1'“" S,a,es P a « id P a *“ » mtemltionaTprogra^e

r b t7”;tnr e ^

pNit,c™'s“sfactfon Urren ' ly ad ° P ' ed ' hrOU 8 hou ' ,he world - a " d *>7 '» maximise global

S: n i 0 r UniCatl0n . S Infrastructure - C1 ^ar and established means of communication
between the co-operating parties are essential. Such means need not be extremelv formal
(indeed, the models show us that the best communication is often informaT) bufthev
must be distinguishable and active. This mandate goes for all levels of co-operaflve

engaged in7e %£*** *”** * ^ “ P^ners, and ~

3.1.1.2 Optimisation Means

Once these constraints have been fulfilled, thus enabling the viability of the project, there
remain some parameters that will help maximise its likelihood of success. Being aware of
^ 0 rHffiml tips due to inappropriate political environments [Section

t .he° establishment P of an international fratneworh

within which Ra is to be achieved.

p m nhasts on an hr ternational Co-operative Nature - The benefits brought by
international co-operation have several origins.

The first one has already been mentioned and is economic: given the limited andoten
Sreas ng financial resources available for space activities throughout the world, he
oSy way we can meet Ra's ambitious scientific and application ob,echves is to share the

resultant cost among several countries.

The second benefit is both technological and scientific: the more participants, the more

i£ — FaS: progress^ Legible
end products and benefits with all the associated risk sharing.

The third is political: if the prime rationale for Ra remains science and its applications we

consider thit the political improvements in international relations it can bring jS part 0

the Success. 3 criteria. A successful co-operation within Ra would be beneficial, since it

would strengthen relations among numerous countries, among which some hard y
would strengtnen re & and hopefully co i our these relations with

friendship In the era of globalisation, we want Ra to help efforts toward global peace^
Also Ra's political and managerial success would be beneficial in being an example and a
modef iXr^erco-opeJon in other areas such as medical research, environment

protection, or industrial development.

a Tjcp nf Fxistine Assets — In some crucial areas, the background level of

expertise requisite for the sSccesT^f Ra is still fragile. For instance, from the scientific
e r^: \o,ar science is relatively young, and from the political and managerial one
^^“emtive ZL has never been attempted. Therefore, we need

to use the assets available world-wide as much as possible.

The practical consequences of this are twofold. First, it means optimising all the national
resources: scientific?technological, human, financial, legal, political, and geographic. W
. 1 t r trade-offs among these different resources, for example technologica ,

financial and political, and consider them unavoidable^ev^heles^the overal
optimisation is certainly one of the parameters that will determine Ra s success.

Co^rmHlv it means building Ra's international framework preferably based on current
S Zu Y International bodies require a long time before having enough proficiency and
m ° t rm to be effective and the newly formed are sometimes received suspiciously.
Thev can be represented as heavy objects moving in a viscous medium: momentum is
7 ariH we more easily change their direction than set them in motion.

tool when establishing Ra's international

framework.

.. . r^mnlpYitv _ The more complex a mechanism, the more likely the

dysfunctional modes“Hus is well known for engineering designs, and it also holds true
t Y r management structures, where dysfunctioning means for instance making bad or no
decirionsf wasting time and money, and favouring inter-personal clashes. Therefore, we

first need to avoid any extra layer of bureaucracy in the decision making process, just as
an architect hunts for sophisticated non-necessary devices, and secondly to keep sound
overall success oriented priorities while allocating tasks.

A1 1° in thls d ° main ' t ra de-offs — if not incompatibilities — among national expectations
optimisation of resources and global efficiency will be unavoidable; but here is also one of
the challenges Ra is willing to address: building an efficient, yet mindful of all
international co-operative framework.

j yhnimum Vulnerability — The strength of a chain is no more than that of the weakest
link, which means that vulnerability has to be assessed for each participating country
agency and even company, and at every level: political, financial, technological, scientific,
human, etc. We will not address in detail each of the latter in this section, but rather
emphasise that Ra's framework would be better chosen keeping the following questions
in the background of considerations:

• Are the participants likely to have a long term local political and financial
support?

• Is there a way to increase this likelihood (if necessary)?

• In case of withdrawal, what back-up solution can be implemented, how fast
and at what cost?

A good example of the kind of decisions political vulnerability considerations can drive
has been described in sections 2.2.9 "Emergence of Smallsat Technology" and 6.11.5
Future Opportunities", which deals with the spacecraft configuration choice. We advise
a fleet of small, almost identical spacecraft, each of them being entrusted to a country or
agency as far as design and integration are concerned, with a possible constraint to use a
commonly designed bus. We thus:

• facilitate national or agency approvals.

• facilitate the overall management.

• reduce unwanted technology transfer.

• make a "reasonable" use of inter-dependence.

• reduce the consequences of withdrawal.

We consider that these factors will contribute to a more favourable and stable political
and managerial framework. r

3.1.2 Developing a Model for International Solar Exploration and
Applications

Co-operation in space is by no means a new phenomenon. Spacefaring nations have
engaged in co-operative activity since the inception of spaceflight; indeed the first
satellites were launched as part of an international collaborative effort known as the
International Geophysical Year. The purpose of this section is to provide an overview of
some available examples of international co-operation in space, and to draw upon the
lessons learned from these examples, both positive and problematic, in developing a
model for international co-operation in solar exploration and applications. We have used
the categorisation of positive and problematic here for the purpose of simplification and
ease of reading. However, we do not intend to imply that it is a matter of taking past
experiences all or nothing into consideration for Ra. No single model can be said to fully
contain all the good or bad experiences from which we can draw. Later in this report, we

Political & Economic Environment • 19

will recommend exactly how these lessons can provide the foundation upon which we
can build an international co-operative framework to fully implement Ra's strategic plan.

3.1.2.1 Problematic Examples

Europa — The Europa launcher provides us with a good example of a programme that
failed mostly for managerial and organisational reasons. It is good to keep it in mind
while trying to set up an appropriate international framework for Ra, so that we do not
repeat the same destructive mistakes [de Dalmau, 1996].

European co-operation in space dates back in 1960, when the United Kingdom was
searching for international co-operation to support its "Blue Streak" endeavour. It was
soon followed by the signature of the European Launcher Development Organisation
(referred to as ELDO) convention by governments of UK, France, Germany, Italy, the
Netherlands, Belgium and Australia, in 1962.

ELDO was to develop the three stage-launcher Europa, whose breakdown method
consisted in chopping the rocket up into almost autonomous parts, then entrusted to the
participating governments. UK would provide the first stage, France the second,
Germany the third, Italy would take care of the payload, Belgium the tracking, the
Netherlands the telemetry, and launches would take place from Australia (later from
French Guiana).

The programme had to face three series of difficulties:

• Economic first, beginning in 1964, when the cost estimates doubled and
later quadrupled.

• Political then, from 1966 to 1971, with the withdrawal of UK from the
programme.

• Finally technical, as of 1967, with a number of failures.

Europa did not manage to survive them and the programme was cancelled in 1972,
without any payload delivered into orbit.

The lessons learned from this sad story can be summarised in the main factors that led to
the failure:

1 . From the beginning, a political top-down approach was mostly carried out.
there was no prime contractor, governments kept financial and decision
power on what was done nationally, ELDO had very limited authority. For
example, ministerial conferences had to be organised for every important
decision.

2. No initial mission, clear responsibilities, rights and management method
had been defined.

3. The political motivations were very different from one country to another:
UK wanted to prove that it was a reliable partner to join the European
Community, France was seeking access to British technology.

4. The levels of development of rocket technology were also quite different.

5. All lacked experience in such a multi-national project.

In conclusion, the whole project failed due to inappropriate initial institutional decisions
and lack of experience. It is also worth noting that the lessons learned from it have
helped in the success of the subsequent European launcher: Ariane.

J pternational Solar Polar Mission (Ulysses) — The National Aeronautics and Space
Administration and the European Space Agency signed a Memorandum of
(MOU ) in 1978 to co-operate on an International Solar Polar Mission
( ). e agreement was for each agency to build a single spacecraft for solar

exploration. The European probe was to fly by the Sun's North pole, while the American
craft was to fly over the Sun's South pole in a co-ordinated, simultaneous trajectory. Both
spacecraft were to be launched on the same Shuttle flight, and the United States would
provide the nuclear power source for the ESA spacecraft, as well as the spacecraft support
m flight through the Deep Space Network (DSN). The intended launch was 1983
[Johnson-Freese, 1990].

The sequence of events that subsequently transpired with respect to the American
contribution to the ISPM is now widely acknowledged as a painful, but valuable, learning
experience for the European Space Agency.

The ISPM was one of five new start requests in NASA's budget for Fiscal Year (FY) 1979.
The mind set was premised on an expanding NASA budget in the out years to
accommodate the maturation of all programmes. However, President Ronald Reagan's
Administration planned a series of domestic civilian spending cuts in its first term in
office. These plans led to a domino effect that, when coupled with increasing Space
Shuttle development costs, ultimately led NASA to cancel the construction of the
American ISPM spacecraft. It was the manner in which the matter was handled,
however, that places the ISPM here as a problematic example of international space
science co-operation. 1 r

NASA exhibited a surprising (to Europe, that is) lack of political will when it came to
defending the ISPM. In 1981 European Space Agency (ESA) Director-General, Erik
Quirstgaard, was notified of the NASA intention to cancel its ISPM spacecraft only a few
hours prior to the Reagan Administration's announcement of budget cutbacks. The
amended FY 1982 U.S. Federal Budget allowed for only US$584 million for space science
as opposed to the previously intended amount of $757 million. This large budget cutback
was the impetus upon which NASA predicated the necessity of cutting the funds for an
entire spacecraft outright. While NASA was admittedly beset by a variety of constraints
which arguably made the spacecraft cancellation a necessity, the attitude which NASA
relayed about the position that American actions put ESA in was not a very sincerely
sympathetic one. The fact that NASA's withdrawal jeopardised the European investment
in ISPM went almost unacknowledged. The lack of consultation by NASA with ESA
prior to the decision was the primary cause of ESA's tension. The decision taken was a
unilateral one, without any real consideration given to alternatives raised by ESA. In
short, while ESA thought it understood the precarious nature of the American budget
process, it at least felt it could count on NASA to fight for what it had committed itself to
in an MoU. When NASA failed to do so, ESA was left not only with a single ISPM
spacecraft, but a bitter uncertainty about America's reliability as a partner in space efforts
[Johnson-Freese, 1990].

NASA did intend to continue to support its contributions to the European spacecraft, including the
radio-isotope thermal generator, the American experiments, the use of the DSN, and the launch
aboard the Shuttle, although the last committment would have to be delayed until 1986.

2 For a concise, but detailed political history of the cancellation of the U.S. ISPM spacecraft see
Johnson-Freese, 1990.

In summary, the ISPM is a problematic example of an international co-operative effort
because it:

1. allowed the withdrawal of one partner to jeopardise the entire mission.

2. was premised on an incomplete understanding of obligations and interests.

3. lacked clear lines of communication.

4. had generated insubstantial domestic political support and will.

5. involved extremely substantial sums of money, and therefore consisted of
large portions of the involved agencies' science budgets (related to 4).

International Space Station — As an ongoing project, the International Space Station (ISS)
is a well known example of international co-operation. While ISS has been successful so
far in co-ordinating the efforts of all partners involved (The United States, Russia, ESA,
Japan, and Canada), its turbulent history has some valuable lessons of which Ra is taking

note.

Begun in 1984 after U.S. President Ronald Reagan invited the American "friends and
allies" to participate in the development and operations of an orbiting space station, what
is now known as ISS has undergone numerous redesigns and adjustments for a variety of
reasons. Several "descoping" redesigns due to American budget constraints were only
the beginning of an extended space station history that always seemed to have an
uncertain future and a delay in development. In addition, the space station project has
repeated many of the same mistakes made during the ISPM. The high political visibility
of the space station, however, has given it its own set of advantages and disadvantages as
an international co-operative effort.

Space station has seen the same American propensity for unilateral decision making as
experienced under the ISPM. When the Russians were brought into the collaborative
effort, it was done so without consultation with the European, Japanese, and Canadian
partners in the venture. The deal was presented fait accompli once NASA had issued the
invitation to Russia.

While the invitation to Russia highlighted an undesirable American decision-makmg
methodology, it did provide the U.S. political system with a more sustainable rationale
for the ISS. Since Russia joined the project, domestic American political support for ISS
has wavered little. When President Reagan called for a space station with allied
participants in 1984, the initiative was an artefact of the Cold War between the East and
West. Upon the dissolution of the Soviet Union, space station supporters attempted to
transfer its justification to science. In a time of diminishing U.S. budgets for space,
however, the American space science community fractured and support for space station
was not’forthcoming. Bringing the Russians in provided an overarching political
rationale, stabilising station's political support. By engaging the Russian space
community in station work, the U.S. had a powerful incentive with which to persuade
Russia to comply with agreements such as the Missile Technology Control Regime, a
political objective much more central to American domestic and foreign policy than an
orbital station for science. Conversely, while the marriage between Space Station
Freedom and Mir II (to form ISS) had the effect of bolstering political support in the
United States for the project, in Europe it served to emphasise the unilateral mind set of
the Americans toward the endeavour, effectively endangering European political will for

the effort.

The decision-making mechanism of the American space complex notwithstanding, the
sheer size of ISS (and associated costs), coupled with its origins, has made it exceptionally

... e t0 dom - hc P olltlcal considerations. One year before inviting the Russians in
the station survived a vote for cancellation in the U.S. House of Representatives by a
mg e vo e. One year later, after Russia joined the programme (and provided the
orementioned rationale), the station survived a similar motion by a margin nearing one

1995RAM- . InE , mope : ESA ' s commitment to ISS was not finalised until the October
n . 5 L ESA Ministerial meeting, the outcome of which, just a couple of months earlier, had
not been assured. Even more recently, budget constraints made Canada seriously
consider withdrawing from the ISS; only after extensive consultations with NASA did
anada commit itself to building the ISS remote manipulator arm. 3 Only in Japan has the

oZ^aTer nafi on r er WaVered ' bUdf!e,ary a " d P °' iKcal fl >Ktuafions in the

What lessons can Ra learn from the experience of the International
most pertinent can be summarised as follows:

Space Station? The

1. Political attention to projects is proportional to their size. The higher the
mterest, the more likely that the project is subject to changing domestic
political winds. Meanwhile, positive aspects of this attention can be high
level political support, but changing domestic political environments can
en anger this. Additionally, the higher the political interest, the more
likely it becomes that "micro-management" by political figures and/or
bodies hinders the project.

2. There must be a sustaining rationale for any space project.

3. International co-operation can be a sustaining rationale (especially
concerning Russia at this point in time).

4. Internal, cohesive scientific community support is essential to domestic
political will (if rationale is closely tied to scientific return).

5 reliability 1 dedsi ° n making harms partner trust and alters perceptions of

Additionally, ISS has re-emphasised the lessons learned from ISPM There is
considerable danger involved in projects where the withdrawal of one partner can

IsTssentfaWn ^ hwestiDait of the P artaer ™«<>ns. Open communications

is essential to good will between partners, and may help alleviate tensions, especially

when dealing with the American budgetary process. ^ ^

h should be noted that what has not remained the same between ISS and ISPM is the
political will on the part of NASA as an agency with regards to the project. Contrary to
ESA s experience with ISPM, at the highest levels of NASA ISS has always been top

priority. Regardless of the internal reasons for this, it is a precedent for international co¬
operation that should be emulated.

3.1.2.2 Positive Examples

Co mmittee for Earth Observing Satellites fCEOS) — Founded in 1984 on the
recommendation of the Economic Summit of Industrialised Nations (G-7), CEOS is an
inter-governmental, inter-agency committee intended to serve as the focus of
international Earth observation co-ordination. The committee now has a small budget for
a secretariat, but there is no permanent staff. In this manner the involved nations keep an
informal structure, avoiding a more rigid and bureaucratic organisational form [U.S.

3 Originally, Canada was to contribute the arm plus the "hand." Under the new arrangement NASA
wi buy the hand and Canada will sustain its committment to provide the robotic arm.

Congress, Office of Technology Assessment, 1993]. 4 Agency representatives meet to
discuss current and future Earth observation systems and their related issues such as data
dissemination and compatibility. Representatives then bring back to their home
organisations information on world plans for Earth remote sensing.

While CEOS itself has no decision-making powers, the information exchange that it
enables, coupled with its forum for policy and engineering issues, has resulted in effective
world-wide Earth observation co-ordination. CEOS recommendations are taken seriously
in the member agencies, and where possible (the most usual restriction being funds),
follow-through has been clearly evident.

In summary, CEOS is a positive example of an international co-operative framework
because it:

1. minimises complexity;

2. possesses high domestic political will (it sprung from high political levels
— very good for sustainable support);

3. Is based on an internationally recognised immediate need;

4. has a clear, definable rationale;

5. has enabled nations to contribute to the international Earth observation
programme in a flexible, independent manner;

6 maintains an informal organisational structure that empowers national
agencies instead of divesting them of power, thereby generating
bureaucratic incentive for the agencies to follow CEOS recommendations.

Tntpr- A vencv Consultative Group for Space Science (IACG) — The Inter-Agency
Consultative Group has been, arguably, the most successful example of international
space science co-operation of the space era. Begun in 1981 on the primary initiatives of
Roald Sagdeev, Director of the Soviet Space Research Institute, and E.A. T * e "delenburg,
ESA Director of Scientific Programmes, the IACG membership consisted of NASA, ESA,
ISAS and IKI (USSR) [Johnson-Freese, 1990]. Its purpose was to co-ordinate the
numerous Halley's comet flyby missions in 1986. NASA was the only agency involved
that did not have a dedicated Halley's comet spacecraft, but it did contribute substantially
bv tracking the crafts with the DSN. The co-ordination consisted of arranging
complementary trajectories, instrumentation, and rapid data evaluation and turn around.
The organisational structure proved so successful that the ad-hoc IACG became a
permanent organisation in 1985 with the purpose to,

"maximise opportunities for multi-lateral scientific co-ordination among
approved space science missions in areas of mutual interest. The lACt, is a
multi-agency international forum in which space science activities are discussed
on an informal basis among representatives of member agencies [Johnson
Freese, 1990].

In the terms of reference it is specifically stated that the IACG does not have a formal
planning role for future missions, nor is it intended to supplant bilateral co-operative
efforts. In fact, many of the IACG's "Core Missions" are bilateral efforts that operate as
well within the IACG framework [IACG, WWW].

The IACG's second project has been the co-ordination of the International Solar
Terrestrial Physics Programme (ISTP). Begun in 1986, the Cluster satellite constellation

< Member agencies are: CNES (France), CSA (Canada), CSIRO (Australia), DARA (Germany), ESA
and Eumetsat (Europe), INPE (Brazil), ISRO (India), STA (Japan), NASA and NOAA (U.S A.) <md the
Swedish National Space Board (Sweden) [U.S. Congress, Office of Technology Assessment 1993].

was to mark the final ISTP core mission launch. Currently, the IACG is working on

choosing lts third project, the first steps for which will most likely be taken in the
December, 1996 meeting [Huber, 1996], 7

The IACG has three ongoing working groups:

Science working group (WG 1): Approximately three scientists from each

member agency participate. This group works to define co-ordinated
science objectives;

• Data exchange working group (WG 2): Co-ordinates the data needs of WG

and has established an IACG Science Information System. Membership
consists of mvolved agency and community scientists;

• Mission design and planning working group (WG 3): The planning that this
group does is co-ordinating trajectory changes, etc. to maximise the science
return for WG 1., not any future mission planning [IACG, WWW],

Additionally, the IACG forms ad-hoc panels to study areas of space science that would be
suitable for future multi-lateral co-ordination efforts. Currently, there are three panels:

• Very Long Base Interferometry (Panel 1)

• Planetary and Primitive Bodies (Panel 2)

• High Energy Astrophysics (Panel 3) [IACG, WWW].

It is interesting to note that although the IACG does not have mission planning powers

ThefirT 3nd rT,H eCtS haVe beneflted from a de S ree of international joint planning.

infprilf Pn ?K Ct/ Ha ey r S COD f *' Was a uni( I ue event that generated widespread scientific
interest, and because of its relatively long-period, the 1986 flyby was viewed as a unique

opportunity to study a comet. The second project, ISTP, was predated by a significant

a Tr T 0]0mt T C f science P lannin g on *e parts of NASA and ESA [Johnson-Freese,
1992J. This joint planning led to other partners being pulled in throueh bilateral
agreements (such as ISAS and IKI). Thus the existence of complementary misLns on the
agendas of national space agencies was not coincidence. While the IACG does not have a
future mission planning function, its projects have been shaped by joint planning prior to
the engagement of the IACG co-ordination role. F

becau^it haS SUrViV6d 3nd successful] y co-ordinated international space science return

1. has limited itself to a single project at a time;

2. keeps itself small at the decision-making level. Around 1990, when the
involved ISTP missions began to number more than twenty, the IACG
began to exhibit the bureaucratic inefficiency typical of a large multi-lateral
organisation. There was co-ordinated movement among the longer-
standing IACG scientists to get the group back to the small, informal
structure that it had under the Halley's comet phase. The move has

^I*® ol ! dated internal IACG support and helped it to more clearly define its
ISTP objectives;

3. does not have a functional allocation role, and is therefore not a threatening
body to the organisational existence of the member agencies;

4. is a grass roots organisation with cohesive support among the space science

community. r

International Mars Exploration Working Group (IMEWG) - Put together in 1993, the
International Mars Exploration Working Group has the mandate to:

• "Produce an international strategy for the exploration of Mars beyond the
currently approved missions (emphasis added).

• Provide a forum for the co-ordination of future Mars exploration missions.

• Examine the possibilities for an International Mars Network Mission as the
next step beyond the 1996 launch opportunity ["Together", 1994].

Agency representatives meet every six months to discuss co-ordination issues of planned
missions. In 1994 the working group submitted to the Committee on Space Research
(COSPAR) a strategic plan for Mars Exploration with international participants. At t e
same time they also published their findings as to an International Mars Network
Mission. The former findings have resonated more fully with national space agency
objectives than the network mission. The network might involve ceding some decision
making power to a central body (such as a multi-national scientific committee),
something that the national space agencies are not yet prepared to do.

The IMEWG seems to have a mixed record of success. While it has fulfilled its mandate,
its relative programmatic influence is debatable; the network mission has received litt e
serious mission consideration. On the other hand, the working group is a useful forum
for the exchange of Mars plans, allowing interested parties to come together in more
discrete co-operative ventures. Furthermore, it does form a venue for the co-ordination of
current and planned missions, allowing agencies to both avoid unnecessary duplication
and increase their relative scientific return from individual missions.

The IMEWG was caught in an unfortunate situation; its submission of findings came
during a period of decreasing national space budgets (with the notable exception of
Tapan) Accordingly, agencies with diminishing budgets approach new mission
proposals warily, and if the mission involves the concession of decision making power
(something that the shrinking budget is already draining) then the organisational
incentive to pursue that option is minimal. Conversely, during times of increasing funds,
the concession of decision making power over a portion of the budget (that grows
smaller, not larger, with time) is less threatening to the agency as an institution.

The limited success of the IMEWG can be attributed to:

1. a clear, definable rationale. While "Mars exploration" is vague in scientific
terms, it points definitively to spacecraft on and around Mars and is
something to which the public can easily relate;

2. its versatility. The continued existence of the IMEWG is not tied to a
specific mission, but to a long term goal. Therefore setbacks, such as
mission cancellations or payload descopings, do not lead to the dissolution
of the working group, allowing its advocacy of Mars exploration to

continue;

2. its established communications infrastructure;

3. an informal, grass-roots nature.

5 Member agencies of the IMEWG are: ASA (Austria), ASI (Italy), BNSC '-(Vf-)' 2 ™S (P™**)' c ~SA
(Canada), DARA (Germany), ESA (Europe), IKI (Russia), ISAS (Japan), NASA (U.S.A.) [ Together

1994"].

S~ Iy ' i! he internatlonal s P ace exploration community was not ready to concede

real tt° n 7. m 8 ° rder t0 make an Int «national Mars Network Mission a

reahty. It » reasonable to assume, then, that in the intervening two years as funds have

become even more restricted for national space endeavours, that the Ituation remains the
lesson ta^nd ° n framework for *>Iar exploraHon and applicaHons must keep this

3.1.2.3 Summary of Lessons Learned

The lessons learned from past and current examples of international collaboration in
space activities served the Ra team as valuable guides in the creahon of an intemlhonS

indtv d rah | Ve framcwor f for solar exploration and applicaHons. While the presentaHon of
md v dual examples above summarised the lessons extracted from each protet h is

ana[ysis° "" 6re * he m ° St im P ortant fusions the Ra team has drawn from this

Political scientific, and technological objecHves of the partners need not be
identical, but they must be compatible.

P ° SSeSS 3 mUtUal understa ndmg of all parties' obligations

• Clear and open communications structures are essential.

® UStai " able rationale is integral to generating the requisite
domestic political will for international co-operation.

Internal, cohesive scientific community support is a prerequisite to
generating domestic political will if the sustaining rationale for the project
rests heavily on scientific objectives. P ]

Flexible and versatile contribution structures enhance the viability of
in ernational co-operative efforts by minimising their vulnerability to
programmatic alterations and political whims. y

• The withdrawal of a single partner from the effort should not structurally

cripple the entire effort. y

• both'fL° St n ; issions/ jP acecraft increase a projects chances of acceptance at

both the national and international level. V

• Unilateral decision making is to be avoided as a barrier to building trust

between partner nations. 6

Advisory international bodies are more acceptable to national bodies than
those with functional allocation roles.

• An internationally recognised, immediate need is a good basis for

international collaboration. 8 S *° r

3.2 The Working Group on International Solar Exploration and
Applications (WG ISEA) F

A successful international framework should allow that the programmatic means bv
the'lACG STeMEWG ‘ ^ - We ^

., LG , the IMEWG. Particular mission cancellations/failures do not jeopardise
the entire return of the venture. This philosophy coincides very well with today's move
oward smaller spacecraft and more focused mission by mission objectives. The flexible
ramework permits large, multi-lateral collaboration efforts while allowing naHnnc i

forTrafJsT^Th Uild ^ indiv j dual s P acecraft (or form their own bi-lateraUgreements
afts). This is a particular concern when we consider the domestic 8 political

implications of transferring industrial opportunities outside the na«on te„us^ of
international co-operation; it is a politically volatile issue at best. Small, discrete cr
brine less attention from domestic economic interests and are more easily defended
against charges of "job exportation." In turn, small craft lead us naturally to multi-lateral,
flexible co-ordination and planning mechanisms, such as that proposed here unde-: the
name, the "Working Group on International Solar Exploration and Applications, or WG

ISEA.

This section will overview the WG ISEA purpose, representation, and structure. In
conclusion, we will evaluate how the WG ISEA meets the constraints on in^ emational
frameworks as outlined in Section 3.1.11 and the optimisation means of Section 31.1.2.
Implicit in our formation of the WG ISEA and its structural evaluation is an awareness of
the lessons learned from our review of international space co-operation models.

3.2.1 Purpose

The WG ISEA should serve as a forum for economical and innovative solar and
heliospheric mission planning, co-ordination, and implementation. In this <™deavmir, the
WG ISEA should take into full consideration the specific mtemahonmeeds of bo* he
space science and space applications communities [Section 5.1 and 5.2). To do this,

WG ISEA should have three functions:

• Strategic Framework planning for national and international solar and
heliospheric science and applications initiatives (the findings for which
should be submitted to the agencies involved) 6 ,

• Current and planned mission co-ordination,

• Reference model design for standardised spacecraft bus configuration (cost
driver).

3.2.2 Representation

To best fulfil its purpose, membership in the WG ISEA should not be limited to the
exploration agencies alone. Here we have detailed a three-tier representation system
which includes Members, Adjunct Members, and Advisors. Full membership should
include any national or multi-national agency which contributes either:

• a primary spacecraft for the programmes we involved in the Strategic
Framework,

• a significant portion of a primary spacecraft, defined here as forty percent,

• an equivalent capability such as deep space tracking.

The space science community has seen these criteria successfully used to designate full
membership in the IACG. However, the WG ISEA goes beyond the IACG in that it is
open to any*organisation that contributes, not just space science and exgorationagencies.
Some potential applications-oriented members include NASDA, NOAA, and UbA .
Chairs and other important offices would rotate between these first-tier members.

6 Applications initiatives fall within the venue of the WG ISEA insofar as they are information-
gathering and/or developmental in nature. Once a system goes operational, then responsibility for its
coordination, etc. would naturally be moved beyond the WG ISEA framework. See section 3.4 on
models for solar warning and forecasting organizations. _

1 TT_C, QrionrP and HllTTianitV

"l 001 !? ti6r f members (ad i unct members) include those bodies which contribute
ei er hardware for mission spacecraft, principal and co-investigators, or adjunct ground
support or observations. Potential members here are national space science and
applications agencies, solar observatories, and universities.

In addition, advisory status should also extend to:

• Mr!!H al s P ace applications advisory and implementation bodies (CSW
NSWC),

• international space applications working groups and services (ISES),

• spacecraft manufacturers (including university small satellite programmes),

• current and potential private users of space environment warning and
forecasting systems,

potential industrial interests in financing an eventual private space
environment warning/forecasting system.

With respect to the international and national advisory bodies and services, there is likely
o e significant overlap in membership, since international representatives are usually
persons involved in national efforts. ^

Obviously, there is a danger of having too many participants creating disparate objectives
and conflicting methodologies, thereby failing to minimise complexity. In order to better
manage this risk, we have proposed a three tier membership (Members, Adjunct

n, e ri erS ' and Advisors ) t0 better delineate the structure. The Ra team believes, however
that the integration of basic and applied science (seeking knowledge for knowledge's sake
vs. see mg knowledge for a pre-designated end) requires the active involvement of the
app icatiom community in the planning process. Thus, membership must extend beyond
the traditional national space research and development agencies.

Members are distinguished from adjunct members in the WG ISEA by both their
contribution and by the allocation of decision-making powers within the group. Advisors
are those groups whose expertise and input to the process are valuable, both to the
working group and to the advisors themselves; they participate on the invitation of the
majority of Members. More detail will be given on this structure in the following section.

.2.3 Structure

part of working
group

Members

Advisors

Organigram of the Working Group on International Solar Exploration and
Applications.

In figure 3.1 solid lines represent direct or adjunct membership status in the working
group. Dashed lines represent advisory and consultative relationships. Dr. Martin Huber
of ESA will ask the IACG to form a Panel on solar investigations this December at the
1996 IACG Plenary session [Huber, 1996]. If this is approved, interaction between the
Panel and the WG ISEA should help delineate scientific objectives and areas of consensus.
This relationship is likely to be informal, as the same scientific community involved on
the IACG Panel will also be key to the WG ISEA. However, it is important to note e
advisory relationship because the responsibilities of the WG ISEA extend beyond basic
science, and thus decisions will incorporate views that transcend IACG considerations.

We would like to note here that the IACG is not an appropriate vehicle to co-ordinate
solar exploration and applications for a variety of reasons.

• Advance planning of future missions is necessary. Currently, agencies do
not have solar missions budgeted.

• Integration with the applications community is advisable; therefore
membership should include applications agencies to best incorporate their

needs.

• Work on a standard reference model spacecraft bus could not take place
within the current IACG framework.

• The nature of the IACG, a very successful organisation as it is, would be
permanently altered by an expansion to include the above groups.

Instead, the Ra team has decided to model the WG ISEA as what could be described as a
ynthesis of what we needed from the structures of both the IACG and the International
Mars Exploration Working Group. Additionally, the WG ISEA has some innovative
eatures such as applications-onented memberships and the reference model sub-
workmg group that should enhance its capability to fulfil its objectives. By keeping the

minimispf^ nd CO '°J' dinatlon ^ t the working group level, bureaucratic inefficiencies are
lmised as much as possible. Additionally, the working group structure provides a
exible mechanism through which outside expertise can more easily be accessed a
particularly necessary capability when considering private sector participation.

3.2.3.1 Standardisation Design & Consultation Group

This support working group is the forum for the creation of a standard international
rence model spacecraft bus for inner solar system exploration. A considerable cost
mi« er in SI1 i a 7 ^ Crete missions is design of the spacecraft. Like robotic planetary
bTcairrih 0 a f7T ° ra u° n 3nd observation lends itself well to international co-ordination

^US Cres^ 19931 acc0I " plis 1 hed | with ^all, relatively low-cost platforms

[ n ' Crest# 1993] ' k an ‘nternational solar and heliospheric exploration initiative nations
can save money while still building their own, discrete craft because the objective^ o^each
craft are limited and the basic bus design standardised. This support group will bring the

Th°p ' 7° invo vement of the private sector spacecraft manufacturers in the WG ISEA
The interaction of science, applications, and the private sector will ensure the

SibiliW Thif ^ pUrp ° Se ° f th<? WG ISEA ' as wel1 as its economic

easibility. This support working group should take into full consideration the

recommendations and opportunities outlined in Section 6.11 of Ra on spacecraft
modularity and standardisation. spacecraft

3.2.3.2 Implementation and Co-ordination

This support working group is responsible for mission co-ordination, data dissemination

m1ss7onT thTt er have b COnS,deraMOnS n ^ ‘° Supp ° rt ,he solar and heliospheric

missions that have been approved in the member agencies, maximising the

complementary data return to the fullest possible extent. This group should also co!

ordinate mission data with Earth-based complementary data sets and historical

observations of interest to the science and applications communities.

3.2.3.3 Strategic Framework Mission Planning Group

The Strategic Framework and Mission Planning Group evaluates and recommends the
Strategic Framework and its constituent missions to the member agencies. This goes
beyond what a panel on the IACG may do. This group is responsible for developing 8 ^
r \ 57r e r° rk f ( wh »ch we recommend closely mirrors the framework outlined in
fhP m k integrates and looks beyond currently planned and/or studied missions in
member agencies. This includes both science and applications missions of the
member agencies as well as integrated efforts (science and applications instrumentation

Add> e Sa ,T7 f U 1S technical] y fusible and scientifically appropriate)

Additionally, the planning group has the task to incorporate the scientific priorities of the
applications sector into the overall mission planning process. That is, the applications
community has demonstrated particular information needs concerning basic solar
science. Therefore, the prioritisation process for scientific missions should incorporate

e^~*Xr S : h 15 m thlS manner SdenCe and a PP lkati0ns ca " most

3.2.3.4 Funding Support

Although the WG ISEA does not develop or operate solar exploration and application
missions directly, costs are associated with the meetings that the WG ISEA will holdand
with the reference design projects the WG ISEA undertakes. It is suggested that funds for
WG ISEA overhead and reference design projects come from first-tier members, perhaps
with membership contingent on contributions to cover these minor costs.

3.2.4 Evaluation of the WG ISEA According to Our Constraints and
Optimisation Means

The Ra team has attempted to fully incorporate the lessons learned from past co-operative
experiences in order to build an organisation that best meets our catena for■«««*“?
international framework, the WG ISEA. In this section we assess the WG ISEA against
our original criteria in order to assure that the working group may function successfully
within the current environment and resources for international co-opera ion m so a
exploration and applications.

3.2.4.1 Constraints

Identifiable, Sustainable Rationale — The rationale of the WG ISEA, to "serve as a forum
for economical and innovative solar and heliospheric mission planning and co¬
ordination," can be argued to be both clearly identifiable and yet imprecise. A rationale
such as, "to land a man on the Moon before the turn of the decade, is much mor
definitive However, the difference lies in whether the same rationale is sustainable.
There may need to be a trade off between precision on the level of the working group s
stated purpose, and sustainability. If the WG ISEA's goal were to implement a then b
then "c" - in that order - then the project (and therefore the WG ISEA) would be
unsustainable if funding for "a" was not available. The issue comes back to the ability of
people outside the community to identify with the rationale of the working group and its
efforts In this respect, the WG ISEA is the facilitator of national interests and the
national interests of the involved parties will transmit their own, individual, yet
complementary rationales to their publics.

S ufficient Domestic Political Will - To argue for science and long-term economic
payback as the basis of domestic political will is tenuous at best. Rather, it is the
economical and innovative nature of the WG ISEA's efforts that will form the backbone of
domestic political support. Today's space participants have general domestic support for
space programmes. The difficulty often comes when justifying the enormous expenditure
of single projects. The WG ISEA and its efforts can generate domestic political will in two
wavs First, the individual mission profiles should be kept relatively small such that they
are not targets for budget cutters. Secondly, their economical and innovative approach to
space activities should draw positive attention as a means to continue national efforts in
space without the exorbitant price tag.

r nhPsive Scientific Community Support - The WG ISEA should take into consideration
both the solar and heliospheric scientific communities. Both specialities supply valuable
knowledge for potential explanations and predictions of occurrences such as coronal
mass ejections and solar flares. In doing this, the WG ISEA avoids the pitfall of splitting
scientific community support when there is no substantive reason for doing so.

Identifiable Funding Sources — Some of the missions outlined in Ra have the potential to
be included in currently funded space agency programmes. The NASA Discovery
programme, for instance, could potentially include some of the small spacecraft missions
included in Ra. In ESA, the Horizons 2000+ indicates its M4 mission simply as Solar

0.1 • Ra- The Sun for Science and Humanity

5*2” ex P loration " Additionally, the C5 Mercury mission has the potential to serve as a
nrn m « some s,m P^ e solar observation instruments |Scoon, 19961 Overall the
uTSIm, .T” , Pr ° POSed by Ra S * ar,S Smal1 and builds •» the long te° it key
do s ve y r C ril; ,hat ° r « aniSin * on “ ™lfi-la,eS

missions or as a mission “group/' such asKA^s SdafwS

,Z n y nS' any eCOn °™ c rel r iS a ■*»*» —pt. However, L involvement of
• ^ a 1 ? 115 communi ty has indirect economic implications Reliable soace

wh^TX^r^ 8 PUrS ? ed f0r a varie,y °‘ raHon a>“. a significant one of
rlicf u ° sate ^ lte operators and users every year due to solar event

PecHon^Ta^^lfsT^ 8 ”? 1 !;^ a ' S ° “Shandy affected by solar events
l ection 4.3.3.2 and 4.5.3.3J). The potential economic return is very real, then.

O pen Communications Infrastructure — The structure of the WG ISEA is conducive to
good communications^ With established meeting intervals, a horizontal organisation and

thf,A“Sa e WG BEA h0Pe,U ' ly WU1 tbe eommunicafiom^UTOess'of

3.2.4.2 Optimisation Means

St? the* WG*ISeI ~ V ™ le ‘"‘national co-operafion is not the

reason tor the WG ISEA (the reason being the maximisation of return on global solar

science and applications resources), the international co-operative nature aste learned

rom the International Space Station model, is a political positive. However the WG

(similar to^^IAC^and the^IMEWG/ 11 ^^ 0 ^ ,h " S im P™ es «" * he ~"“1p.

manufacturers, industrial interests, and private users is because the
their respective experiences have yet to be maximised by bringing them all Ser
single forum for solar and heliospheric investigations. The combination of perspectives
the potential to be a powerful tool for economical and innovative space activities.

C ° m P leX ^ ,- ^he flip side of bringing all these different experiences together
t the complexity level of the working group is increased. Admittedly, the WG^SEA
s more complex than both the IACG and the IMEWG. CEOS, however, operates at a
omplexity level more consistent with the level that the WG ISEA will experience As
mentioned previously we have attempted to minimise the impact of this complexity on
the working group s efforts by creating a two-tiered membership system and^voidine
hierarchical decision-making. Avoiding the pitfalls of bureaucratic inefficiency is a chief
chaflenge faced by all multi-lateral frameworks. The structure we have given* the^G
ISEA here should aid them in this labour. 8 U

Mi nimum Vulnerability — In one way or another, all the above constraints and
optimisation means all have a single goal in mind: to minimise vulnerability Among

other things, the structure of the WG ISEA is not organisationally threatening to its
participant agencies; it should create bureaucratic incentive for active ^volvemen
through its generation of domestic political will; it is programmatically flexible,
individual investment risk is minimised with the discrete spacecraft approach;

complexity has been minimised as much as possible; clear and team

exist and it makes appropriate use of existing resources and expertise. In all, the Ra team
has attempted to insure the WG ISEA as fully as possible against domestic politics,
individua/mission difficulties, internal inefficiencies, and national funding constraints, to

simply name a few.

3.2.5 Final Recommendations for the WG ISEA

In conclusion, the Ra team believes that the WG ISEA format is the most innovative and
efficient mechanism to pursue international space efforts in solar exploration an

applications.

The Ra report is uniquely timed to take advantage of current space science trends. While
the IACG will most likely form a Panel on solar investigations in the coming year, NASA
is planning to bring its Sun-Earth Connections Roadmap to the American space science
community for assessment. That meeting is set for August, 1997 at Woods Hole [Sun-
Earth Connections, WWW]. It would be very expedient if an international mechanism
such as the WG ISEA could be in place before then. With such a vehicle established, e
NASA Space Science strategic plan could more fully incorporate ^ ter ^°^
opportunities and capabilities into its programmatic recommendations^ Furthermore for
planning to be effective it must take place before the budgetary cycles for the target years
are locked in. Therefore the medium-term scenarios that require an advance planning
element need to be planned in the near future if they are to fly before the year 2010.

Additionally, the culmination of the International Solar Terrestrial Physics Programme
provides a natural foundation upon which to build Ra's Strategic Framework. The near
Sun and heliosphere environment is a natural extension of solar-terrestrial interactions. It
is important, then, to have a timely formation of the WG ISEA to take full advantage of
the current period in space science.

3.3 Data Dissemination Principles for the Ra Project

An efficient and workable scientific data sharing system is one of the most important
issues for the successful operation of the Working Group on Int f J atl °/! al Solar
Exploration and Applications (WG ISEA), as described in Section 3.2. Within this group
a special support group is considered for data dissemination, mission co-ordination, and

operational tasks [Section 3.2.3.2].

A data collection and dissemination system is characterised by a large amount of complex
data from different missions and the various geographical locations of the data users.

In previous years, each solar system exploration mission had its own information
distribution policy and structures. In 1981 the Inter-Agency Consultative Group for
Space Science (IACG) was created to co-ordinate scientific research of the four space
agencies to study Comet Halley [Section 3 . 122 ]. This Group now "provides the means
for the optimally co-ordinated operations and a mechanism for data sharing and join
data analysis"[IACG, WWW].

For the newly developed WG ISEA, special information structures are needed as well as
development of data acquisition, archiving, distribution and exchange principles. Four

models are examined for the WG ISEA. The principles of data dissemination for the
Cluster project is an example of data organisation at the mission level. The International
Solar Terrestrial Physics Programme (ISTP) and the IACG Science Information System
building principles are compared as examples of data co-ordination of multiple missions.
Despite the fact that the organisational structure of the IACG is not completely
appropriate for the WG ISEA working scheme, its "Rules of the Road" for data exchange
seem to be well developed and are also reviewed.

Data distribution issues have three aspects: economic, political and technical. Before
establishing rules for data exchange it is necessary to create hardware structure and
software for data receiving, processing and exchange.

3.3.1 Cluster Data Technical Details and Infrastructure

Forty four instruments from the four Cluster satellites were planned for data collection.
The Cluster Science Data System (CSDS) was established for data management. This
system has the following components:

• National Data Centres (six national centres were directly involved in the
CSDS - Austrian, French, German, Scandinavian, UK, Hungarian, and two
centres in the USA and China were registered via directly involved centres).

• Operations Control Centre

• Joint Science Operations Centre (JSOC)

• Ground-based programmes

• network infrastructure within the CSDS

• Cluster user interfaces

Each Data Centre processed the raw data from a specific set of experiments and makes it
available via the network for the other Centres. The European Space Operations Centre
(ESOC) (Darmstadt, Germany) was responsible for mission planning, data disposition
and bulk distribution of the raw data. ESA provided infrastructure for data and
information exchange within the CSDS and also for data ingestion by National Data
Centres and data manipulation by scientific users.

The CSDS net interconnected the CSDS National DATA Centres, the JSOC and various
ESA establishments. The principal CSDS net is based on the existing ESA infrastructure
implemented as a self-contained logical system from an addressing, routing and security
point of view. The net has been designed to provide a logical interconnection between
local area networks (LAN) across a wide area network (WAN) infrastructure.

The CSDS user interface was developed on the base of the software available at the
European Space Research Institute (ESRIN), ESA's establishment in Frascati, Italy. Under
the overall responsibilities of the Cluster project, ESRIN has tailored the existing software
to the specific Cluster requirements.

A Cluster-specific standard for data exchange based on the Common Data Format (CDF)
was established.

A large community of users with varying levels of familiarity with data manipulation
caused the need to have both a convenient user interface and a solid and reliable network
infrastructure. Given the different configurations, existing at the various national Data
Centres, two versions of the CSDS user interface were developed, one running on Solaris
and the other Open Visual Machine System [Drigani, 1995].

2 - IT.

: ^ it.

3.3.2 Data Dissemination Policy Issues (Cluster, ISTP, IACG)

Policy issues at either mission level or intra-agency level should define common
components: categories of the users, their rights to access the information, types of data,
rules of distribution, periodicity of exchange, types of missions from which data are going
to be analysed simultaneously, services suggested by national data centres, and time
period after which whole data is released to the general public.

Categories of users:

Cluster — principal investigators (PI) and co-investigators (Co-I), general public.

ISTP — principal investigators, co-investigators, associates, students, guest investigators
and general public.

IACG-IACG Community members — principal investigators and co-investigators; non-
IACG scientists who received general approval from IACG; general public.

Established classification of the users defines their different rights to access the
information. Usually the PI and Co-I have raw data or high resolution data access and
can share information from the instruments for which they are responsible with anyone
they choose but are not allowed to distribute another investigators data [Green, 1996].

ISTP and NASA have following data policy rules: NASA-funded missions and
instruments have an open data policy; key parameter digital data will be non-proprietary
and publicly available to identify possible scientifically interesting events or intervals (the
key parameters are not intended for journal publication unless certified by the PI); all
teams will contribute relevant digital process data for any special events or intervals that
are selected for study by the Global Geospace Science (GGS) Team and for the IACG
Campaigns in a timely and responsive fashion; during a validation period of up to three
months after data acquisition, all use of data for scientific investigations must be
approved by the PI whose data are being used; and higher level process instruments and
theory data products, along with their associated documentation and all relevance
software, will be publicly available immediately after validation [Data Policy for
ISTP/NASA Funded Missions and Instruments homepage, ISTP, WWW].

The Cluster data distribution policy is as follows: data such as a summary of the
parameter database and summary data plots have unrestricted access, however the prime
parameter database has restricted access limited to the Cluster community only. The high
resolution data will be handled by the Principal Investigators. The Pis will also respond
to requests for the data from the user community. The CSDS infrastructure will probably
be used to route the requests from the high resolution data.

The raw data would be distributed on a set of CD-ROMs to each participating institute
(about 80 world wide) on a weekly basis to reduce network loading; the network will
contain quick-look data. Data Centres are responsible for the registration of scientific
users, assignment of data access rights and checking of these rights when they access the
data. Not all the Data Centres will offer the same services on-line.

Specific IACG Information System Policy Issues:

During the 1990s, the IACG plans three scientific campaigns, each of which addresses a
set of specific questions related to the solar-terrestrial environment. The scientific aim of
the first IACG campaign is a multi-mission (Geotail, Interball, Wind) collaboration which
greatly extends the interchange of data within the international research community. For

'iz • Rr The* ^nn fnr Sripnrp and Humanity

the first IACG campaign, special "Rules of the Road" were developed for different

,°£ ei ?f-° n »- S Wlt r ^ ata exchange from the spacecraft and ground-base facilities for
identification of the obligations of researchers about data provided by other science
research investigators. "Rules of the Road" consider mission rules, key parameter
distribution campaign rules, IACG membership, membership for the non-IACG mission
scientists, sharing of data issues data set preparations, authorship and public release of
a. Rules of Road of the first IACG campaign are adopted now as rules for all
campaigns. However, each campaign can develop its "special rules" that apply only at
agreed upon campaign times [Green, 1996]. J y

M ission rules — Key parameter data are generated by the IACG core missions and other
ancillary data sources in a common format. During a campaign the key parameter data
C f Te 0115810113 are free] y exchanged and accessible to all principal investigators
(PI) and co-investigators (Co-I). Key parameter data will be used for the multispacecraft
event identification and are not publishable unless explicit certification is given bv the
appropriate instrument PI. b 7

^ A aio P ai ,^ n rules "Rules of the road" govern access to and use of data contributed to the
IACG first campaign database and data analysis. During campaigns, any data base can
be created and included into the information system. "Rules of the road" are mandatory
for all participants and those applying to participate. Even if the member withdraws
from the IACG the campaign is obligated to continue to respect the rules established.

Sharing of key parameter data — Data are routinely exchanged between campaign
members and used to support the identification of events. Members share high level
campaign data products with members of their research team, but are not entitled to
urther distribute the campaign data provided by other investigators. Distribution of
detailed instrument data is the responsibility of the instrument principal investigator.

^ >a ! a Tc e A t c Preparatl0nS — Location of databases is preferably in the centres such as NASA
and ISAS. Access to the database and support software will be provided by individual
members of the campaign.

A uthorship “ When an investigator's data is used in the analysis of an event the
investigator who provided these data should be kept informed of what they are being
used for, should be invited at an early and appropriate time to participate in the
correlative analysis and would normally have the option of being a co-author of any
resulting publication or presentation, including abstracts.

Public release of data — Unrestricted access to a database will be granted at the
conclusion of the campaign; usually the period between proprietary data and open data is
typically 2 years. It is important to note the effect of NASA's new open data policy on the
IACG "Rules of the Road." [Green, 1996] NASA no longer temporarily restricts data
access to mission scientists whereas other space agencies still maintain restrictions for a
certain time period to incentivize researcher involvement. Provisions in the IACG "Rules
of the Road were necessary for non-NASA IACG investigators to be given the
opportunity to withdraw their data from the IACG first campaign database before NASA
public release. [IACG homepage, IACG WWW],

The idea of restricted access has two side effects. From one side it allows the owners of
the instruments to generate new scientific ideas. From another it makes the number of

scientists who work with these data much smaller.

3.3.3 Principles of the Development for the Ra Data Systems

Data dissemination principles within the WG ISEA will take guide-lines from the data
policy of the IACG Science Information system development.

One of the driving ideas behind the WG ISEA is to close the gap between pure science
and applications. Classification of users is suggested to be more specific: basic scientists
such as principal investigators and co-investigators, applied scientists, associates,
students, guest investigators and private operators. Private operators are those who own,
for example, satellite constellations and will be aware of the space weather. Private users
are important for financial reasons. However, consideration of restricted access to the
raw and high resolution data over a two year period is important.

In order to minimise costs, the Ra team suggests utilising for data purposes, system
structures that have already been developed by different national agencies for solar
science missions (structures for SOHO, Cluster, Interball, Geotail). Data processing
application needs and distribution to the private users must be added to the main
national data centres.

It is important to develop the following: formats for data exchange between different
space agencies, types of data dissemination, categories of users, periodicity of data
distribution, and a list of services provided by different centres.

In summary, the main significant features of the Ra user interface and infrastructure
should be:

• minimum training needed to use it (user friendly)

• participation of interdisciplinary and various federal agencies

• international participation and data exchange

• free and open access for the general public to secondary data [Scoon, 1996].

3.4 Organising for Solar Warning and Forecasting

Increased understanding of solar processes and improved technologies for solar
observation present the opportunity to mitigate or prevent damage to human activities
and assets from dangerous solar phenomena, both in the space environment and on
Earth. The creation of solar warning and forecasting services, however, relies on more
than scientific and technical knowledge. It requires an efficient organisation that is
appropriate to the service's resources, tasks, users and political environment. This section
reviews and proposes criteria by which models for solar warning and forecasting
organisations can be judged and introduces current and future models for these
organisations. This sets the stage for the construction of a specific organisational model,
the Inter-agency/International Interface ("Triple I") Model for modem solar warning and
forecast services.

342 Basic Criteria for Examining Solar Warning and Forecasting
Organisational Models

Sections 3.4.2 and 3.4.3 discuss current and future models for solar warning and
forecasting organisations. Before examining these models in detail, it is important to set
criteria to compare and contrast them. This section presents ten general criteria by which
the nine models in sections 3.4.2 and 3.4.3 will be judged for solar warning and
forecasting organisational recommendations in Section 3.5.

3.4.1.1 Adequate Development Funding and Stable Operations Funding

Arguably the single most important criterion for any solar warning and forecasting
organisation is its ability to garner the initial funds needed to erect the infrastructure
(satellites, ground instruments, tracking stations, data archives, data dissemination
networks, etc.) for its solar warning and forecasting service. Some organisations, such as
the U.S. Air Force, already have many of these elements in place and would need less in
the way of infrastructure development than, say, a commercial solar warning and
forecasting service. Likewise, some organisations may have more ready access to
development funds than other organisations. This requisite may prove to be the most
important criterion in coming decades for solar warning and forecasting organisations as
improved services will be dependent on investment in technology, hardware and
knowledge, especially the deployment of in situ instruments between the Earth and Sun
and refinement of various operational space physics models.

After the erection of the service's infrastructure, any solar warning and forecasting
organisation will require a long-term and stable source of funds for operations to ensure
the continued existence of the service. Continued funding will also be important for
infrastructure upgrades and replacements. Long-term funding will thus need to be
flexible enough to accommodate cyclical highs and lows in equipment acquisition.
Certain organisations, like NASA by its own admission with the Space Shuttle, may be

unsuited to operations and reluctant to undertake such funding.

The source of these development and operations funds is also a major consideration for a
solar warning and forecasting organisation because it determines to whom the
organisation is primarily responsible. Civil government, military, commercial and
international solar warning services all have different potential sources of funding that
determine their prime users and political masters and thus the makeup of their services.

3.4.1.2 Take Advantage of Current Solar Warning and Forecasting Capabilities

Future solar warning and forecasting services, if rationally constructed, will take
advantage of current solar warning and forecasting capabilities rather than rebuilding the
necessary infrastructure and reconstructing the necessary knowledge base from the
ground up. This criterion drives intra-governmental (agency level) solar warning
organisations to co-operate with other solar warning organisations within the same
government. It drives governments with solar warning capabilities to co-operate with
each other as well.

The advantage of existing capabilities is an especially important criterion for commercial
solar warning organisations. Without the withdrawal or metering of current government
solar warning services, commercial organisations will find it extremely hard or even
impossible to compete with what are “free” public solar warning services (although taxes
obviously do support such services). Commercial solar warning organisations are more
likely to find market niches or horizontal interfaces within the more comprehensive solar
warning services that governments provide.

3.4.1.3 Simple Structure With Clear Functional Allocation

Because several different organisations currently provide solar warning services, future
services may well be provided by conglomerates of today's organisations. It will be
vitally important to these future services to clearly delineate different functions between
their constituent organisations to avoid confusion, duplication and plays for power. This
does not necessarily require a standing, overarching manager for the whole service but
simply requires thoughtful planning in its organisation. The correct organisational

structure will be simple, clearly outline the functions of each element and grant each
element autonomy in achieving its functions while co-ordinating with the other elements.

One important functional allocation decision will be the division of development (design,
fabrication and launch of the space segment, for example) from operations (spacecraft
monitoring and data acquisition). Another rational organisational division might also be
set at the boundary between raw data acquisition versus data interpretation,
phenomenon modelling, and warning and forecast dissemination. However, if in situ
instruments are placed between the Earth and Sun for solar warning, traditional
functional allocations based on government agency domain over the terrestrial versus
space environment systems may be blurred. NASA and other national space agencies
will likely no longer be the sole operators of spacecraft in the deep space environment.

3.4.1.4 Identified Users

Any solar warning and forecasting organisation will need to clearly identify the users of
its services so it can adequately and reliably meet their needs. These users can be
classified into user communities based on their common backgrounds (civil, military or
commercial) and into user groups based on the commonality between their resources
(satellites versus power grids). The following paragraphs list the possible user groups for
a solar warning and forecasting service and briefly delineate their unique needs (user
groups derived from Space Weather Prediction homepage, American Geophysical Union
[WWW]; Lund Space Weather and AI Centre homepage, Lund University [WWW] and
Spacecraft Anomalies Due to the Radiation Environment in Space homepage, NASA
[WWW]).

Commercial, Civilian and Military Satellite Operators — Solar phenomena can affect
satellites in four ways: heavy energetic particles can penetrate electronic components and
create errors in instrument data or false spacecraft commands, energetic electrons can
shorten component lifetime through dielectric charging, less energetic particles can cause
surface charging problems, and geomagnetic storms can heat and expand the Earth's
upper atmosphere which creates drag on satellite orbits. Satellite operators require
advance warnings of large energetic particle emissions from the Sun (such as flares) from
in situ plasma devices. Additionally, monitoring or modelling of the magnetopause
during a geomagnetic storm is required for geosynchronous satellites to predict when
these spacecraft may pass through the magnetopause boundary and be subjected to
quickly reversed magnetic fields. These quick field reversals can cause dangerous electric
discharges and disorient satellites that rely on magnetic torquing for attitude control.

Humans in Space (Astronauts, Future Employees and Tourists) — Energetic protons from
intense solar flares and large CMEs (Coronal Mass Ejections) can increase the radiation
dosage for humans in space by magnitudes of order in a very short time frame (tens of
minutes from a solar event). Although present systems do provide adequate solar flare
warnings for short stays or small numbers of persons in low LEO orbits, better CME
tracking is needed to ensure safety levels for long duration spaceflight and the predicted
large numbers of future space workers and tourists. Future manned missions to the
Moon, the asteroids or Mars will also require the expansion of current CME tracking
capabilities to new regions of the solar system and better long term predictions of solar
activity (over periods of years) for mission planning purposes.

Civilian and Military Radio Communications Users — High frequency terrestrial radio
waves that rely on ionospheric reflection for propagation near and across the Earth's
polar regions can be interrupted by solar induced ionospheric disturbances. Satellite
radio waves that must penetrate the ionosphere are also altered by these disturbances.
Although television and commercial radio signals can be affected, critical rescue and

military communications are the most vulnerable users. Better accuracy in solar
warnings through in situ magnetometer and better forecasting models, especially the
interaction of the ionosphere with geomagnetic storms, will allow the users of these
critical systems to better predict when they need to seek other means of communication.

Civilian and Military Navigation System Operators — Ionospheric disturbances induced
by solar phenomena in the magnetosphere can alter the path of navigation signals that
transverse the ionosphere (through refraction) or propagate via ionospheric reflection (by
changing the altitude of the ionosphere). Like radio communications users, better upper
atmospheric models that interface with current ionospheric and magnetospheric imaging
instruments and magnetometers are needed to enable navigation system operators to
predict these signal path deviations and correct for them.

Commercial Electric Power Companies — Geomagnetic storms can create disturbances in
the Earth's magnetic field which induce currents (Geomagnetically Induced Currents or
GICs) in long power lines. These currents can destroy transformers, cause generator
heating, and create rapidly and widely varying power levels in transmission lines.
Although power network damage from geomagnetic storms can be measured in the
millions of dollars and is well recorded, techniques to mitigate this damage are poorly
understood and underemployed. Additional accuracy in geomagnetic storm warnings
and forecasts will give power companies the confidence they need to develop, deploy and
utilise adequate GIC countermeasures [Bolduc to Sillen, 1996].

Pipeline Managers — To prevent corrosion in today's buried pipelines, managers pass
small currents through their pipelines to eliminate anode junctions with moist soil. GICs
in pipelines can temporarily negate or even reverse the benefits of pipeline currents.
Pipeline manager requirements are similar to those of electric companies; additional solar
forecasting accuracy is needed to enable countermeasure development.

Industries Using Extremely High Quality Control Manufacturing Processes — Peaks in
the number of control problems in extremely high quality manufacturing processes (those
that limit defective sub-units to a few parts per million such as semiconductor
manufacturing) have been statistically linked to geomagnetic storms, but the physical
connection between storms and the lowered quality in various manufacturing processes
has not been determined. Industrial manufacturers require research on this connection
before they become future users of solar warning and forecasting systems.

Geodetic Surveyors — Surveyors that use the Earth's magnetic field to make
measurements have been long-time users of solar forecasting data. Solar warnings and
forecasts enable surveyors to know when their data is inaccurate due to solar phenomena.
Although their needs can be better met through continued refining of current solar
warning systems and forecasting models, geodetic surveying imposes no remarkable
requirements on future solar warning and forecast services.

3.4.1.5 Capability of Users to Protect Their Resources from Dangerous Solar
Phenomena

Although the previous section identified eight potential user groups and three user
communities for a solar warning and forecast service, the services that such an
organisation provides will be relatively useless unless most of the noted user groups can
protect their resources from dangerous solar phenomena. For example, a solar warning
will not actually protect terrestrial electric power distribution grids from a geomagnetic
storm unless the companies that operate those grids have procedures and equipment in
place beforehand to protect their resources from the storm. Likewise, satellites that rely
on a solar warning and forecasting service must be designed with various active and

Political & Economic Environment • 41

passive countermeasures in mind to prevent damage to the satellites from dangerous
solar phenomena, regardless of any solar warning or forecast. Current user capabilities in
these countermeasure areas are very limited, and the potential countermeasures
themselves are often system specific. Thus the link between a solar warning and
forecasting organisation and its users must also include the technical analysis of user
countermeasures to dangerous solar phenomena. This will require yet another specific
functional allocation within the solar warning and forecasting organisation or require a
third party to perform the analysis needed to erect the physical and procedural solar
countermeasures.

3.4.1.6 Ability to Satisfy User Data Needs

Section 3.4.1.4 classified solar warning and forecast users based on their common
resources (user groups) and on their common backgrounds (user communities). These
differences must be taken into consideration when considering the data needs of specific
users. Some possible differences in data needs between various user groups and user
communities include:

Relevance of Solar Data Supplied to the User (Which Solar Phenomena Are Being
Observed?) — Any solar warning and forecasting organisation will need to concentrate
its observations on those solar phenomena which affect its users. Differences in the solar
phenomena that various users are interested in falls along user group lines because of the
similarities of user group resources. For example, civil, military and commercial satellite
operators will all be interested in the interaction of geomagnetic storms with the
ionosphere while power companies will be interested in interactions between
geomagnetic storms and the Earth's magnetic field. Although the details are technical,
some solar warning and forecasting organisations are better suited to satisfying certain
user groups data needs because they concentrate their observations on certain
phenomena.

Timeliness of Solar Data Supplied to the User (How Often are Solar Forecasts Updated?)
— Different organisations provide different update rates for solar forecasts, and these
differences lend themselves to various user communities which require a shorter or
longer duration between updates. The military user community may require very rapid
updates during times of conflict, whereas the commercial user community's forecasts can
be updated at more regular intervals.

Lead Time of Solar Data Over Phenomena (Does the User Have Enough Time to Protect
His Resources After a Solar Warning?) — Different user groups may require more or less
lead time in order to enact countermeasures to protect their resources. For example,
powering down an electric grid may take less time that reorienting a satellite before a
geomagnetic storm. This criterion will be especially important in the near future as
forecasts and countermeasures are tested and refined through experiential contact with
dangerous solar phenomena.

Comprehensibility of the Solar Data (Can the User Understand the Significance of a Solar
Warning or Forecast?) — Different user groups and communities will possess different
levels of technical knowledge regarding the interpretation of the implications of a solar
warning or forecast for their resources. For example, power companies are unlikely to
have ready access to solar physicists whereas satellite operators may have implicit
knowledge about the effect of solar phenomena on their systems from designing those
systems. Warnings and forecasts will need to be tailored to the technical sophistication of
the user either through the primary solar warning and forecast organisation or through
secondary organisations who take raw data from the primary organisation and interpret
it for different users.

A9 • TVip fnr ^ripnrp anrl T-fiimanitv

3.4.1.7 Warning Versus Forecasting

Until this section, solar warning and solar forecasting have been discussed as a single
organisational service and function. Solar warnings, however, require a level of technical
understanding that falls below that required for solar forecasting. Solar forecasting
requires an interface with human expertise that solar warning does not necessarily
require except in its development phases. Certain organisations, because they already
possess this technical expertise, will thus be better suited to solar forecasting in addition
to solar warning than other organisations.

3.4.1.8 Reliability and Accuracy of Solar Warnings and Forecasts

Although an obvious point of concern, the reliability and accuracy of solar warnings and
forecasts will be important criterion in choosing between different organisational models
for a solar warning and forecasting service. For example, military users may have solar
warning accuracy requirements that are too costly for a commercial service to provide.
Likewise, a military service may lack the expertise needed to generate a long-term
forecast for a commercial user. It will be easier to match the right service provider to the
right user, rather than forcing the provider to change or improve its data gathering and
interpretation methods or forcing the user to cope with less than ideal data.

3.4.1.9 Stability of Solar Warnings and Forecasts Over Time

Although this criterion is partially addressed in section 3.4.1.1 by continued operations
funding, it is also an especially important criterion when considering a military solar
warning and forecast service. National emergencies may require a military service to
temporarily halt the dissemination of solar data to commercial or civil users. Similarly,
civil or commercial services that serve military users in addition to other users may also
be required to limit their data dissemination in times of emergency. Clear internal
policies that conform to national laws must be in place to anticipate these contingencies if
the line between military and civil/commercial solar warning and forecasting is crossed
by either users or providers.

3.4.1.10 Capacity to Incorporate New Solar Knowledge and Technology

Despite the fact that solar warning and forecasting are relatively undeveloped fields, both
scientifically and technologically, any enduring solar warning and forecasting
organisation will find it vital to possess the capability to integrate new solar models and
new solar observing technologies into its warning and forecasting services. Some
organisations are well suited to perform this continuous development in house whereas
others will need to co-operate with external organisations to transfer this knowledge
because they lack the necessary technical expertise and infrastructure. °

3.4.2 Current Models for Solar Warning and Forecasting Organisations

With these ten criteria in mind, four contemporary models for solar warning and
forecasting can be introduced. These models are drawn from existing organisations that
deal with some aspect of solar warning and forecasting.

3.4.2.1 Single Civilian Agency (NOAA — SEC)

Perhaps the simplest organisational model for a solar warning and forecasting service is
that of the single civilian government functionary. The U.S. National Oceanic and
Atmospheric Administration (NOAA) undertakes solar warning and forecasting duties in
addition to its other terrestrial weather services through its Space Environment Centre
(SEC) located in Boulder, Colorado. The SEC, formerly the Space Environment

Political &■ Frnnnmir FnuimnmDnt • aq

Laboratory (SEL), is one of NOAA's seven National Centers for Environmental
Prediction. NOAA obtains solar warning and forecasting data from its Geostationary
Operational Environmental Satellites (GOES) and its Polar-orbiting Operational
Environmental Satellites (POES). NOAA is responsible for processing this data, analysing
it to create forecasts, and real time "nowcast" warnings oriented to meet the needs of
civilian government and some commercial users [Space Environment Center homepage,
WWW].

A single civilian agency like NOAA has several advantages over other organisations
including stable operations funding, a base of warning and forecasting capabilities on
which to draw, a relatively simple organisational structure, defined user groups, and the
ability to continue warnings and forecasts uninterrupted. NOAA, however, cannot
develop new solar observation technology independently, may lack the ability to create
forecasting models, may not provide data services appropriate to military (and possibly
some commercial) users, and does not currently integrate user countermeasures with its
warnings and forecasts.

3A.2.2 Single Military Functionary (USAF — AFSFC)

The United States Air Force (USAF) undertakes the development of new models for solar
forecasting through its Air Force Space Forecast Center (AFSFC) at Colorado Springs,
Colorado. These models concentrate on near-Earth space and include the Parameterised
Real-time Ionospheric Specification Model (PRISM), the Ionospheric Forecast Model
(IFM), the Magnetospheric Specification and Forecast Model (MSFM), the Solar Wind
Transport code (SWT) and the Interplanetary Shock Propagation Model (ISPM). Except
for PRISM, all these models are still under development and current 24 hour AFSFC
geomagnetic forecasts provide, at best, 44% accuracy [Space Weather Prediction Home
Page, WWW]. The AFSFC also obtains a variety of in situ space environment
measurements through the U.S. Defense Meteorological Satellite Programme (DMSP).

Although a solar warning and forecasting service in a military department is
organisationally as simple as a civilian government functionary and derives many of the
same benefits described in Section 3.4.2.1, it is questionable whether a purely military
organisation could promise to provide uninterrupted solar warnings and forecasts in
times of national emergency or whether military user community requirements match the
requirements of civilian or commercial user communities. It is also interesting to note the
emphasis DoD places on solar forecast model development, which complements the
wider solar and space environment instrument arrays deployed on NOAA's weather
satellites.

3.4.2.3 Inter-agency Functionary (SESC)

The United States has resolved the tension between its military and civilian users by
consolidating NOAA SEC resources and USAF AFSFC resources in the U.S. Space
Environment Services Center in Boulder, Colorado. The SESC is staffed by NOAA
civilians, uniformed NOAA Corps, and USAF personnel. It provides forecasts of solar
and solar induced geomagnetic activity through optical and radio indicators and
geomagnetic indices. These indicators and indices are obtained through ground based
observations of solar flares and solar activity, through particle. X-ray and magnetometer
data from NOAA's GOES satellites, from particle data from NOAA's POES satellite, and
from various data from DoD's DMSP satellite. The SESC provides a single, national point
for space warning and forecast organisation in the United States by drawing on the
resources of government agencies whose individual requirements necessitate a certain
level of resource independence [Space Weather Prediction Homepage, WWW].

The advantages of an inter-agency functionary like the SESC are obvious, especially for
commercial users who can look to one public service for their solar warning and forecast
needs. It is important to realise that the SESC does not programmatically co-ordinate
NOAA and USAF resources and thus does not prevent the duplication of agency
capabilities within the U.S. government. °

3.4.2.4 International Data Collection and Dissemination Service (ISES)

Formerly known as the International Ursigram and World Days Service (IUWDS), the
International Space Environment Service (ISES) provides an international data network
for the acquisition and distribution of solar warning and forecasting data. Supported by
various scientific societies, the ISES collects data from ten Regional Warning Centers
(RWCs) throughout the world [International Space Environment Service homepage,
WWW]. RWCs are nationally supported organisations and primarily serve the needs of
their national users. The data contributions from RWCs to ISES can vary greatly and
include data from such disparate sources as Japan's Geostationary Meteorological
Satellite (GMS) [Space Weather Nowcast abstract homepage, WWW] and Australia's
Radio and Space Services [IPS Radio and Space Services homepage, WWW], The SESC in
Boulder, Colorado acts as the clearing-house for RWC data and serves as the ISES's
World Warning Agency (WWA).

ISES is a valuable glue between the world's various solar warning and forecasting
services, although obviously dominated by SESC's contributions. Its purview, like that of
the SESC, is limited to data co-ordination, and it cannot prevent the duplication of
national resources internationally and is extremely dependent on national resources for
service continuity and improvement. ISES clearly defines the functional boundaries
between development, operations and raw data acquisition at the national level and data
collection and distribution at the international level. ISES may suffer from a clearly
defined set of users but is also considering initiatives to improve forecasts from the point
of view of user end requirements.

3.4.3 Future Models for Solar Warning and Forecasting Organisations

Five possible future models for solar warning and forecast services also exist as national
plans, in current meteorological organisations or as theoretical ideals.

3.4.3.1 True National Functionary (NSWP)

Attempts are underway in the United States to consolidate NOAA, USAF and SESC
resources with other agency resources to create a National Space Weather Programme
(NSWP). In 1993, the U.S. National Science Foundation (NSF) was prompted by the
science community to undertake the improvement of solar forecasting capabilities. The
NSF formed three working groups (Sun / Solar Wind, Magnetosphere, and Ionosphere /
Thermosphere) to address the technical and organisational issues involved. Through the
actions of these working groups and the NSF Office of the Federal Coordinator for
Meterology (OFCM), a Committee for Space Environmental Forecasting (CSEF) was
formed. The CSEF wrote the first drafts of the NWSP Implementation Plan and directed
the formation of a National Space Weather Council (NSWC) and a Committee for Space
Weather (CSW which replaced the CSEF) in late 1994. The NWSP Implementation Plan is
now a living, changing document that is continually refined by the NSWC. The NSWC is
a multi-agency oversight and direction group consisting of representatives from DoD, the
U.S. Department of Commerce (DoC — NOAA's parent department), the U.S.
Department of the Interior (Dol), the U.S. Department of Energy (DoE), NASA and NSF.
These representatives act as spokespersons for their agencies and departments in the
NSWC and address issues of individual agency scope, requirements and resource

commitments. The NSWC ensures that common agency needs are met while securing the
planning, programming and budgeting interests of the agencies involved. By its own
admission, the NSWP does not co-ordinate the engineering aspects of the technical
systems of its constituent agencies and relies upon its users to tailor its solar warning and
forecast products to their needs. The NSWC is overseen by the CSW. An important
element of the interaction between the NSWP Implementation Plan, the NSWC and the
CSW is the use of defined metrics to measure the progress of U.S. solar forecasting
capabilities evolution [National Space Weather Implementation Plan homepage, WWW].

The "overarching goal" of the NSWP "is to achieve an active, synergistic, inter-agency
system to provide timely, accurate, and reliable space weather warnings, observations,
specifications and forecasts within the next ten years." Technical objectives to achieve this
goal include the development of accurate 72 hour solar event forecasting models and 48
hour near Earth space weather forecasting models [National Space Weather
Implementation Plan homepage, WWW].

Each agency involved in the NSWP contributes unique hardware and human resources to
the programme. The USAF, in addition to its current observational and modelling
capabilities as described in section 3.4.2.2, has proposed through its Air Force Phillips
Laboratory a Solar Mass Ejection Imager (SMEI) for 48 hour CME warnings. The SMEI
would fly on a Sun-synchronous polar orbiting satellite [Space Weather Prediction
homepage, WWW]. The USAF might also contribute daily CME warnings through its
Over-The-Horizon (OTH) radar to a future NSWP [OTH Space Weather Forecasts
homepage, WWW].

NASA also promises to contribute critical observation and modelling capabilities to the
NSWP. Real time solar wind data from NASA's WIND spacecraft currently provides a
testing ground for potentially very accurate two hour space environment forecasts from a
spacecraft placed at LI. However, even with adjustments WIND cannot constantly
monitor the solar wind, and NOAA is providing resources to modify NASA's Advanced
Composition Explorer (ACE, to be launched in 1997) for the provision of longer term, real
time solar wind data. NASA is also developing the Quantitative Magnetospheric
Predictions Programme (QMPP) in its Space Physics Division which will relate different
regions of solar induced phenomena through WIND and ACE measurements.

The last contributor to the NSWP is the U.S. National Science Foundation (NSF). Through
its Geospace Environment Modelling (GEM) programme, NSF is developing the
Geospace General Circulation Model (GGCM) which is a modular programme adaptable
to the forecasting needs of various users. GGCM complements NASA's QMPP.

Perhaps the most important aspect of the NSWP Implementation Plan is its recognition of
the eventual need to replace the temporary WIND and ACE spacecraft with dedicated in
situ solar warning spacecraft at Lagrange points or in solar orbit. The ability of the NSWP
to co-ordinate hardware contributions makes it a potential vehicle for the deployment of
these spacecraft. However, the NSWP has yet to seek additional contributions to such an
effort outside the United States.

Although the NSWP organisation is not a simple structure and de-emphasises user end
requirements, it is flexible, maximises the use of current national solar warning and
forecast capabilities, rests solidly on the budgets of its constituent agencies, and has the
capability to improve U.S. solar forecasts and sustain forecasting services over time.

3.4.3.2 National Inter-agency Functionary with Foreign Contributions (NPOESS)

A hybrid of the NSWP model and the ISES model is a national inter-agency functionary
that incorporates hardware contributions from foreign countries. The U.S. National
Polar-Orbiting Operational Environmental Satellite System (NPOESS) is a developing
meteorological system that may demonstrate the theoretical operational feasibility of
foreign contributions to a national interagency solar warning and forecasting service.
NPOESS developed out of studies of the convergence of NOAA and DoD polar orbiting
weather satellite capabilities dating as far back as 1972. Increased Congressional interest
in 1993 led the Vice President to recommend convergence, and a Tri-agency Study Group
under the U.S. Office of Science and Technology Policy (OSTP) was formed in 1994. The
OSTP recommended convergence to the U.S. Congress and the President in 1994. A tri¬
agency ad hoc conversion transition team was established, and in October 1994 the team
established the Integrated Programme Office for NPOESS. In May 1995, a tri-agency
Memorandum of Agreement (MOA) between NOAA, DoD and NASA was signed. In the
MOA, NOAA and DoD agreed to provide a total of $1.4 billion for NPOESS acquisition
through 2001, NOAA became the lead agency for NPOESS execution and operations,
DoD became the lead agency for NPOESS acquisition, NASA became the lead agency for
technology transition, and the involvement of the international community was
recognised.

The NPOESS Integrated Programme Office consists of an Associate Director for
Acquisition from DoD, an Associate Director for Operations from NOAA and an
Associate Director for Technology Transition from NASA who all report to an NPOESS
System Programme Director. A Joint Agency Requirements Group feeds input to the
Associate Directors while a Senior Users Advisory Group confers directly with the
System Programme Director. Above the System Programme Director, an Executive
Committee consisting of the DoD Under Secretary for Acquisition and Technology, the
DoC Under Secretary for Oceans and Atmosphere and the Deputy Administrator of
NASA holds power and is advised by a Joint Agency Requirements Council [Williamson,

In terms of physical hardware, the U.S. portion of NPOESS consists of two common. Sun-
synchronous, polar orbiting weather satellites; one procured with DoD funds and one
procured with NOAA funds. At this level, the NPOESS organisation resembles the solar
warning and forecast capabilities currently shared between NOAA and DoD in the SESC
with additional hardware co-ordination. However, NPOESS also includes a third satellite
contributed by ESA and Eumetsat that carries both European and U.S. instruments.
European participation grew out of NOAA budget overruns, which forced NOAA to look
for partners to take over this responsibility. NOAA and Eumetsat drew up a plan to have
ESA and Eumetsat assume half of NOAA s civilian morning-crossing operational
meteorological data responsibility through Eumetsat's METOP polar satellites. NOAA
found a partner to be responsible for hardware in Europe before political pressure forced
NOAA and DoD to co-operate domestically, and this European partnership was folded
into NOAA and DoD agreements. Further co-operation with the Russian polar orbiting
meteorological satellite, Meteor-3, is also being considered as a serious possibility [U.S.
Congress, 1993].

The direct integration of discrete foreign hardware in a national, interagency co-operative
structure makes NPOESS a unique model for a future solar warning and forecast
organisations beyond the current SESC, NSWP and ISES structures. The NPOESS model
also clearly separates functional responsibilities based on the unique advantages of each
participant. The NPOESS model may be especially applicable when solar warning and
forecast services decide to deploy dedicated solar and space environment observation
satellites at Lagrange points or in solar orbit. The high development cost of such systems

may require burden sharing beyond that which any national, interagency organisation
can provide.

3.4.3.3 Regional Convention Organisation (EUMETSAT)

Although a solar warning and forecasting service is unlikely to be based on a regional
organisation because of the global impact of solar phenomena, the European
Meteorological Satellite Organisation (EUMETSAT) does provide a possible model for
international co-operation in solar warning and forecasting. The convention creating the
EUMETSAT organisation was ratified in June of 1986 for the exploitation of ESA's
Meteorological Satellite Programme or METEOSAT (the first European geostationary
weather satellite had been operational since 1977). EUMETSAT is a classical international
organisation, governed by a Council with representatives from all member states for issue
arbitration and resolution. The day to day functioning of EUMETSAT is undertaken by a
small Director's secretariat. Although ESA is still charged with the development and
launching of new METEOSATs and the European Satellite Operations Center (ESOC)
handled the data acquisition and daily operation of the METEOSATs until 1995 (both of
these functions are arranged in a separate agreement between ESA and EUMETSAT),
EUMETSAT is responsible for METEOSAT administration and financing. METEOSAT
financing is accomplished through mandatory contributions from signatories to the
EUMETSAT Convention. If contributions are withheld, EUMETSAT data is not provided
to the signatory in question. It is important to note that EUMETSAT, ESA and ESOC do
not analyse METEOSAT data. That function is instead carried out by national
meteorological agencies which are signatories of the Convention and by the European
Centre for Midterm Weather Forecasting [van Traa-Engelman, 1993].

Future international solar warning and forecasting services might wish to utilise aspects
of the EUMETSAT organisation, namely the consolidation of administrative and financial
functions under an international management. This international management overlay
stabilises funding, allows for national processing of the international data stream, clearly
delineates functional boundaries and provides a vehicle for data and hardware co¬
ordination. The two inappropriate aspects of the EUMETSAT organisation for an
interagency or international solar warning and forecast service are (1) the integration of
resources on single spacecraft designs and (2) the nature of EUMETSAT data release,
which is dependent on participant contribution. These aspects of the EUMETSAT
organisation are made possible by the increasing interdependence and unification of
European states but would probably not be possible in a rival interagency setting or a
global international setting.

3.4.3.4 True International Functionary

Given enough time, an international agency under the aegis of the United Nations or a
service funded through similar national contributions might possibly emerge as the
world provider of solar warning and forecasting data. However, the need for solar
warning and forecast data is not yet great enough to warrant the expenditure of limited
international resources on such a service and an international agency would likely still be
extremely dependent on national solar warning and forecasting resources, limiting its
independent yet international character. International data collection and dissemination
services like ISES are more likely to continue as the primary means of international co¬
operation in solar warning and forecasting. If international co-operation in solar warning
and forecasting does extend beyond mere data co-ordination into hardware contributions,
then the Eumetsat model (with services dependent on treaty membership and payments
supporting the hardware) or the NPOESS model (independent but co-ordinated
hardware contributions) will probably emerge well before any true international solar
warning and forecast agency.

48 • Ra: The Sun for Science and Humanity

3.4.3.5 Commercial Service

It is theoretically possible that a commercial entity could undertake all the functions
necessary to provide a solar warning and forecasting service. Competition with
government services available to the public makes that possibility unlikely, however,
unless government users are willing to rely upon a commercial provider and unless
governments are willing to eliminate, meter or transfer their solar warning and forecast
services to a commercial entity. Bureaucratic inertia in the case of civilian government
services and security requirement rationales for military services makes both of these
contingencies distant propositions, however. There is also the question of just how
commercial such a service would be since its primary customers would continue to be
government users and because it would likely be a monopoly once established,
preventing the market entry of equal competitors. There may be a market for a
commercial solar warning and forecasting service, but that market can probably
accommodate only one major provider.

In the foreseeable future, the commercial world is more likely to fill horizontal gaps in
government solar warning and forecasting services by adding value to those services
rather than by creating its own vertically integrated service. Some potential gaps for
commercial entities to fill include: the development of countermeasure routines for
specific satellites, power grids and other systems threatened by dangerous solar
phenomena, the real time interpretation of government warnings and forecasts for less
technically literate users, and consulting regarding the impact of solar phenomena on
user resources. An example of value added commercial activity in solar warning and
forecasting is ARINC Incorporated of Colorado Springs, Colorado, which developed a
Space Weather Training Programme for the USAF Space Command and 50th Weather
Squadron and a solar effects flowchart under DoD contract [Davenport, G.R., WWW],

3.5 Recommended Organisational Structure for Future Solar
Warning and Forecast Service Services: The

Interagency/Intemational Interface ("Triple I") Model

Based on the ten criteria for an ideal solar warning and forecast organisation in section
3.4.1, none of the nine current and future solar warning and forecast organisations in
sections 3.4.2 or 3.4.3 address all the possible shortcomings of such organisations. It is
necessary to derive a unique model to approach the ideal match between solar warning
and forecast services and the current political and economic environment in which they
exist.

3.5.1 Themes for the Construction of a Modem Solar Warning and Forecast
Organisational Model

Several themes can be drawn from the critical review of the nine current and future solar
warning and forecasting organisations in sections 3.4.2 and 3.4.3:

1. The United States is by far the predominant actor in solar warning and
forecasting services throughout the world. Actions undertaken by the
United States will critically affect any international solar warning forecast
efforts and must take the international context into consideration.

2. The United States is taking sufficient measures to sustain and enhance
interagency co-operation to reduce the costs of solar warnings and forecasts
and to synergise advances in its total capabilities without endangering the
independence of these individual agency services. The SESC and NSWP
are central to achieving these objectives.

Political Fronomir Fnuimnmonl * AQ

3. The international solar warning and forecasting community possesses an
adequate vehicle for data collection and dissemination in the form of the
ISES.

4. The international solar warning and forecasting community lacks an
organisational means to collectively improve solar forecasting models and
solar warning systems. This is partly because these advances require
national political, military and budgetary commitments and partly because
of the dominant role of the United States.

5. Future advances in solar warning and forecasting will require investments
in two key areas: the refining of forecast models and the deployment of
dedicated in situ solar and space environment instruments outside Earth
orbit. The former is realisable within certain agency or national resources,
but the latter will be highly dependent on resource contributions, risk
sharing and cost sharing between agencies or governments without
threatening the independence and ability of those organisations to meet
their own user needs.

6. Public government organisations are likely to remain the primary providers
of solar warning and forecast services for the foreseeable future.
Commercial services can, however, assume secondary roles left unattended
by government services.

7. Even with greatly improved solar forecasting models and solar warning
systems, a gap may exist between very accurate solar forecasts and the
ability of users to take advantage of a forecast's warnings.

8. Advances in solar warning and forecasting will be highly dependent on the
application of basic research into the Sun and its effect on the space
environment.

3.5.2 Requirements and Structure: Constructing the

Interagency/International Interface ("Triple I") Organisational Model for
Modem Solar Warning and Forecast Services

Taking these eight themes into consideration, it is possible to recommend an
organisational model for future solar warning and forecasting organisations. The
requirements of this model should include:

1. Sustain intra-governmental efforts like SESC and NSWP to co-ordinate,
consolidate and improve national solar warning and forecasting
capabilities, especially space environment modelling.

2. Continue international solar warning and forecasting data collection and
dissemination (ISES).

3. Expand international solar warning and forecasting service co-ordination to
the level of hardware contributions. The NWSP can facilitate this effort by
identifying and involving potential international partners according to the
NPOESS model.

4. Share risks and distribute cost burdens among the number and type of
participants needed to achieve 3.

5. Clearly delineate functions according to the strengths of national and
international participants as in the EUMETSAT and NPOESS models.

6. Maintain an open dialogue with basic solar and heliospheric research
organisations.

50 • Ra: The Sun for Science and Humanity

7. Provide a focus for user end requirements. Commercial solar warning and
forecast services are appropriate for this role.

These requirements lead to the organisational structure shown in figure 3.2.

Fig. 3.2

T» nple *• l y° de J : ^ organigram of evolving solar warning and
forecast organisational relationships emphasising the critical role played
by the interagency and international interface. Note the dashed line
separating development and hardware roles from operations and data

The critical, currently non-existent junction in this structure is the
Interagency/International Interface, and this organisational model is appropriately
named the Triple-I Model for Solar Warning and Forecasting Service Organisation to
emphasise that interface. It is possible that the role of the "Triple-I" box in this
organigram could be filled informally through international NSWP outreach Given the
recommendations in Section 3.2, however, the "Triple-I" function could also be more
formally instituted through the applications side of the proposed Working Group on
International Solar Exploration and Application (WG ISEA).

3.5.3 The "Triple I" Model and Its Relationship to a Proposed Solar Warning
and Forecast Spacecraft Constellation °

In chapter 9, a minimal, mid-term, in situ , solar orbiting constellation of ten to twenty
small spacecraft in the ecliptic, each carrying a magnetometer and a plasma instrument
for space warning and forecast applications is introduced. It is suggested that the "Triple
I model presented here is an ideal model for the development and deployment of such a

Political & Economic Environment • SI

solar warning and forecasting constellation. The mission definition, standards and
reference designs for spacecraft contributions to the constellation proposed in chapter 9
would be developed jointly through the “Triple I" model, but each participant would be
responsible for the actual acquisition, launch and operation of its own spacecraft. Data
sharing would occur through existing channels like ISES in the “Triple I" model.

3.5.4 A Thought for the Future: Will Solar Warning Spacecraft Become the
First Operational Deep Space System?

If a solar warning and forecast organisation, regardless of its makeup, does deploy solar
monitoring spacecraft beyond Earth orbit to protect terrestrial and space based resources
it will likely mark an important transition in human space activities. Although national
space agencies and even military functionaries have undertaken scientific, exploratory
and technology demonstration missions beyond Earth orbit, no organisation has ye
deployed spacecraft beyond Earth orbit for an immediate, "practical, operational
rationale. Many have predicted that the first human robotic activities in deep space
bevond science, exploration and technology demonstration would involve resource
gathering or even colonisation on other celestial bodies. This section predicts, based on
the history of human space activities in Earth orbit which was initiated and dominated by
communications and remote sensing satellites, that the first operational human activities
in deep space will be solar and space environment monitoring spacecraft in solar orbits or
at various Lagrange points. The significance of solar warning and forecasting
organisations will lie not only in the economic benefits that may be derived from their
services, but also in the important historical footnote they will provide as humanity
expands its presence in the universe.

3.6 Solar Research and Forecasting in the Context of Russian
Space Policy

Current Russian space policy was initiated in February 1992 with the foundation of the
Russian Space Agency by a Decree of the President of the Russian Federation.

The Soviet space industry began its development in the late 1950s in the Ministry of
Defence ('Sputnik" was designed as extension of the development of intercontinental
ballistic missiles). During the Soviet era, there were multiple ministries and committees
(such as the Ministry of General Machine Building "Minobshemash", Academy of
Sciences etc.) which were involved in the space industry, but there was no single agency
responsible for space development in general. During the Cold War period, space policy
was aimed at preserving the strategic military balance and political leadership between
USSR the USA and their partners. Changes which occurred in Central and East
European countries in the late 1990s shifted national governments space policy goals
towards broader international co-operation in space exploration, as well as in global
security and environmental problems.

3.6.1 Current Situation

Russia inherited the major part of the Soviet space industrial complex. Since 1991, newly
independent states have started the transition to a free market economy. The transition
period is characterised by an unstable political and economic situation, undefined time
boundaries and an unclear programme of further development (nobody can predict now
what type of society will exist in Russia after the transition period). In such tenuous
times, planning becomes even more difficult but must nevertheless continue.

Russia is aware of the potential developments in the national space industry and has
made the following steps to support national space activities: a) the foundation of the
Russian Space Agency in 1992, b) the resolution in 1992 of the Government for the
development of the Federal Space Programme, c) adoption of the Russian Federation
Space Activity Law in 1993 (now under revision in Parliament) and d) the government
resolution on Space Activity support in 1994 [Mironjuk and Pieson, 1996].

The Russian Space Agency serves both as state customer and the major space technology
manufacturer, providing operation co-ordination for the enterprises and organisations
involved in space activities. The Russian Space Agency is responsible for space policy in
the Russian Federation: 7

• development of the Russian Federal Space Programme

• development of scientific and applied space technology

• co-ordination of scientific and applied commercial space projects

• further development of research and testing facilities in the Russian space
industry

• international co-operation as well as co-operation with CIS states.

The Russian Federal Space Programme, together with the resolutions of the Government
of the Russian Federation, define the development of the space activity. The main goals

of the Russian space policy were formulated by the Russian Federal Space Programme as
follows:

• fundamental and applied space exploration and Earth monitoring;

• use of space industry benefits for the national economy, scientific, technical
and social progress;

• ensuring the Russian Federation defence needs and control of the fulfilment
of the arms control agreements

• international co-operation in the interests of world scientific, technical and
social progress, global environmental monitoring, world space market
participation.

The Russian space industry suffers today from the general tendencies of the current
economic situation in Russia as well as from the specific issues of the legacies of USSR
space policy. Negative issues of the current economic situation in the country include: an
economic crisis and a decrease in industrial production; absence of the well developed
private sector; absence of customers with sufficient funds for the space services inside the
country [Moscow Aviation Institute Space Economics Department, 1995], An
unwillingness of the newly created financial structures to invest money into the state
industry together with high level of militarisation of the space industry; absence of
competition space projects and absence of the independent expertise, make life of the
space industrial enterprises more difficult and complex [Hozin, 1995],

However the Russian space industry, despite all the problems mentioned above, has very
high scientific and technological potential, especially in such fields such as booster
design, telecommunications, navigation, remote sensing, biotechnology, microgravity
materials processing, manned spaceflight and dual use of military technologies.

Commercialisation of the space industry in Russia became one of the important issues in
Russia after 1991. International co-operation and establishment of the new world space
markets are the primary challenge for future development of the Russian space industry.

The Russian Space Agency is aware of the developing domestic space market, as well as
need for participation in international space markets.

Today space commercial activity is controlled by state through licensing of various
activities by the Russian Space Agency. Therefore, search and rescue operations, natural
disaster and emergency warning as well as weather forecasting are excluded from
commercial space activity. The state has exclusive rights to own cosmodomes with all
launching facilities. Foreign investors are allowed now to have not more than 49% in the
property of the joint companies dealing with space activities [Moscow Aviation Institute
Space Economics Department, 1995].

Commercialisation of the Russian space industry is going slowly because of inflexible
structure of the management, decision-making marketing strategies and developed user
infrastructures.

3.6.2 International Co-operation Within the Ra Strategic Framework

The International Co-operation Department of the Russian Space Agency is responsible
for co-operation with other space agencies and organisations. In the later stages of
negotiations, the Office of the Federal Space Programme Planning can be involved to
include future missions into the Federal Space Programme. Usually the institutes of the
Russian Academy of Sciences, such as the Institute of Space Studies ( IKI ), Institute of
Terrestrial Magnetism, Ionosphere and Radio Wave Propagation ("IZMIRAN") are the
principal investigators from the Russian side in solar and interplanetary missions.

Due to its unstable domestic economic situation, Russian participation in current
international space projects have been limited but can take place through the following
channels:

• contributing intellectual property

• provision of a spacecraft bus for a research programme with an
international set of experiments and instruments

• conversion of military technologies or dual use of military technologies

• building space equipment through direct financing by foreign
organisations.

3.6.3 Possible Russian Space Programme Contributions for the Near-Term
RA Strategic Framework

The Russian space programme can suggest for the Near-Term Ra Strategic Framework
the current mission "Interball" as well as different meteorological and military satellites
under conversion which have instruments for measuring geophysical parameters in a
near-Earth orbit which are already functioning or planned to be launched in the
framework of the Space Segment of the Unified State System for Eco-monitoring [Scoon,
1996; Johnson-Freese, 1996]. For example, the meteorological geostationary satellite
"Electron" is part of the Russian meteorological system "Planeta-C." It was launched in
November 1994 and has special instruments for helio-geophysical monitoring on board.
It provides measurements of protons at 0.2-500.0 keV, electrons at 0.2-2.5 MeV, particles
at 2.0-12.0 MeV, UV emission at 10-130 nm and gamma rays at 0.2-1.0 nm. It also
measures variations in the direction of the Earth magnetic field [Zhdanovich, 1994].

The development of the Unified State System for eco-monitoring needs special
consideration. The concept of the Space Segment of that system is based on the
unification and further development of existing Russian remote sensing systems as well

as systems for space weather monitoring into one global informational system with
common control centres, various data analysis centres, and user terminals at different
evels [Bondur, 1995], The Space Monitoring System is based on the multi-level
hierarchical principle with the various spacecraft flying in different orbits, with a wide
range of instruments on board and a network of ground stations. In the framework of
this space segment a few declassified systems are suggested to be utilised:

H -?nn Y .! tem f °:.r an , C °r 1: " Legenda " — S P ace s X stem "Legenda" (circular orbit,
inn in, v ' = 1 } W \ h radar which S ives ima § es 1 00x100 km 2 or strip with the width
H - inn ^ Z r f c °l“ tlon f?-. 1500 m and satellite spacecraft with circular orbit

P . , m and 1 ~ legenda includes "Diagnosis" instruments for the mapping of
the Earth s magnetic field, "Pole" instruments for the forecasting of Earth eruptions, and
Predvestmk instruments for the monitoring of the ionosphere and magnetosphere and
ground stations for the measurement of the electromagnetic fields on the ground.

S ystem for global monitoring: "Oko-1" and "Oko-2" - These spacecraft monitor Earth in
real time They use two types of orbits: geostationary and half-day elliptical. They can be
utilised for understanding the helio-geophysical situation and diagnosing the complex
phenomena of the space environment. Oko includes "Reis" spectrometers for hot and
cold plasma detection; differential proton spectrometers; electron, proton and alpha
particle spectrometers; plasma sondes for the measurement of the velocity and density of
the solar wind [Bondur, 1995]. y

System for space weather m onitoring: "Prognoz" and "Orion" — The system for direct
monitoring of space weather is based on the "Prognoz" satellite. Two satellites of
Prognoz-M (first apogee 20,000 km, second apogee 200,000 km) have two ion and
electron spectrometers and "Reis" instrument complexes. Two more space weather
monitoring spacecraft are also planned: "Orion-C," for the measurements of the
parameters of the near-Earth space different from the direction to the Earth heliocentric
angles and Orion-Sl, planned to be put into a libration point orbit at 1.5 million km.

3.6.4 Russian Space Programme Contributions for the Mid- and Far-Term Ra
Strategic Framework

For the irud-term and far-term, it is possible to use space science experience and research
in I ™ sslon strat egic planning, as well as solar missions which were included in
the Federal Space Programme up to 2000 but are not able to be fulfilled because of the
difficult economic situation in Russia. One of these projects is "Solar Zond" - to studv the
Sun as a star from the distance 5 solar radii. The Russian space industry can provide the
following platforms for future solar missions: space buses and sub-systems for the joint
designs [Pieson, 1996]; launchers such as "Energia", and "Proton" etc, spedal heat
protection materials; and robust engineering. An example of the resources Russia has to

7 no/ r ^ a needl j' sha P ed space probe with a cone looking towards the Sun which reflects
70 /o of the incident photons, allowing only 30% of the them affect the space probe, which
reduces thermal system protection requirement by a factor of three [Marov, 1996].

3.7 International Agreements and Contracts in the Ra Strategic
Framework 6

In order for the Ra programme to advance, co-operation between government bodies and
contracting private companies is required. Section 3.7 reviews the types and forms of
international contracts, involved when co-operation among and between government
bodies and private companies occur. 6

The inter-governmental agreement would also refer to applicable state obligations and
responsibilities found in the United Nations treaties dealing with space law, the Quiet
Space Treaty of 1967, the Liability Convention of 1972 , and the Registration convention of

1975 .

3.7.1 Co-operative Agreements Between Governments

The Ra project involves co-operation between government bodies. This co-operation can
render the following benefits:

• reduction of cost to individual participant countries,

• maximising the potential of achieving programme objectives,

• risk sharing,

• limiting the ceiling of liability,

• increased support base across the national/international spectrum.

There are also disadvantages to involving government bodies in a programme such as Ra,
which must be considered. These include:

• potential funding uncertainties

• lack of coherency and continuity in decision-making processes

• susceptibility to political processes

Although international co-operation has some potential risks, as discussed above, there
are also substantial benefits to be gained. These benefits far outweigh the risks.

3.7.1.1 How to Co-operate

In a co-operative programme of the type proposed, an inter-governmental agreement is
required. An inter-govemmental agreement will include discussions on major items such

as:

• how expenses will be shared

• designations of responsibility for facilities and decision-making

• intellectual property rights

• registration, jurisdiction and control

• ownership of elements and equipment

• proposed design and development timetable

3.7.1.2 Plans for Utilisation

From past examples, however, it is recognised that agreements of this type need to be
flexible. Differing legal requirements among countries dictate the desirability of building
a legal framework which allows individual countries to fulfil their own bureaucratic and
political requirements, and permit the structure to evolve along functional lines that will
best maximise the potential for programme success. A successful example of such an
arrangement is the Tamamushi agreement concluded between ISAS and NASA in 1986
[Johnson-Freese, 1993]. The agreement allowed both agencies to fulfil their bureaucratic
needs while flexibly allowing the programme for which it was created, Solar-A, to
proceed.

3.7.2 International Industrial Contracting

Each country in the world has different domestic laws. Therefore when the government
is contracting with another country's private company, or between other country's
private companies, a detailed, written contract is necessary.

Contracts are routinely concluded between governments and private companies. Some
types of contracts include a fixed cost contract and an upper limit cost contract.

3.7.2.1 Fixed Cost Contracts

Fixed cost contracts decide costs at the beginning of a project. If the conditions of the
contract have not changed, the cost has not changed. But if the conditions of a contract
have changed, the cost changes. When the objects of a contract have a market price or
have been made before, a fixed cost contract is the most economical and simple contract
form. Fixed cost contracts are awarded using a bidding system.

3.7.2.2 Upper Limit Cost Contract

Upper limit cost contracts decide the cost of a project with a rough estimate in the
beginning of the contracting process. After finishing the work, the government bodies
check the actual money spent to fulfil the contract, and the government bodies and
private companies decide the final cost. When the objects of the contract are developing
something new, upper limit cost contracts are the most common contract form. It is
impossible to correctly estimate the development cost of new objects. However, it is
important to set some limits on the cost because the budgets of government bodies are
limited. Additionally, the upper cost limit is a warning against wasting money to private
companies. However, upper limit cost contracts make a lot of work for the government
bodies, limiting the number of upper limit cost contracts used.

3.7.2.3 Intra-Industry Contracts

This type of contract is useful in the case of very big projects, for example, in the case of
developing and making a new satellite in Japan. Company A is the prime contractor,
company B makes antennas, company C makes batteries, and company D makes sensors.
The prime contractor takes the responsibility to fulfil the contract, assuming
responsibility for the work of the subcontractors. This arrangement is easier for the
government bodies because they need to oversee the prime contractor only.

The Ra Strategic Framework includes new, internationally interconnected projects and
brings new factors into consideration. It is important to use different contract types as
designated by the environment and objective of the contract.

3.8 Concluding Remarks

Chapter 3 outlined the political environment in which solar exploration and applications
must take place by examining previous examples of international cooperation in space
science and various organizational models for solar warning and forecast services.
Criteria were introduced and important lessons learned by critically examining the
history of international cooperation in space science and the organizational schemes for
solar warning and forecasting services. Out of these lessons, two critical
recommendations are made. First, those national and international bodies involved in
either solar research or solar warning should form an international Working Group on
International Solar Exploration and Application before August 1997. The second is that
international solar warning and forecasting cooperation should be improved by stressing

Political & Economic Environment • 57

coordination at the Interagency/Intemational Interface, either through the WG ISEA or
through international outreach by the U.S. National Space Weather Program. If these
steps are taken, solar exploration can look forward to a more coherent and sustainable
future, and solar warning services can begin to mount the modelling and spacecraft
infrastructure needed to improve their forecasts.

Thp Sun for Science and Humanity

Chapter 4

Our View of the Sun

Since the earliest day of humankind we have observed the Sun crossing the sky every day
in an apparently never ending cycle. From the worship of ancient cultures to our current
scientific study of the Sun, there has been a great change in the way humans see the Sun
as well as a steady development in knowledge.

The intention of this chapter is to provide the background information for why we study
the Sun, how this study has been attempted throughout history and how solar science is
performed today. Furthermore, it should stress the questions about the Sun that lead to
the objectives given in chapter 5.

The chapter is divided into six sections. The first gives an overview of how ancient
cultures have seen the Sun and leads to the sections where our discovery of the Sun is
described in a modem scientific way. Section 4.2 introduces the Sun as a star and section
4.3 presents the phenomena in interplanetary space. Section 4.4 describes the basic Sun-
Earth interrelations and section 4.5 the effects of the Sun on humans and technologies
Section 4.6 closes this chapter by suggesting how the Sun may be used as a resource.

4.1 Studying our Sun

Our earliest observations of the Sun are reflected in the myths and artefacts of various
cultures, which demonstrate the various levels of sophistication humans have had in
their understanding of the Sun. Modem solar science, however, will touch the mysteries
of the Sun in ways that our ancestors could never have dreamed. But in many ways our
motivation for this exploration remains mythic in nature. We are the first generation that
can undertake this journey through spacecraft-based science. What we discover will
likely change the way we view our solar system, the universe and, ultimately, ourselves.

4.1.1 The Sun in Myth and Legend

To early humans, the Sun was surely one of the most awesome forces in their daily lives
and perhaps the most celebrated. Its power warms the air, grows food and materials for
fuel and shelter, and drives the cycles of wind and rain.

Myths about the Sun are found throughout most cultures. Although these stories vary
greatly, they give a glimpse into the importance of the Sun within various societies. As
Indo-European peoples spread throughout Europe, India, Iran and Asia Minor, they
spread the concept of a high sky god. This sky deity quite often faded in importance
leaving the universe to his offspring, usually the Sun god.

In Africa, it is common for the Supreme Being to be expressed as a Sun god. For example
the San believed that the Sun was once a mortal being who emanated light from his
armpit. Children of the village wanted to make the light brighter so they threw him up
into the sky where he still shines now as a round disk for all mankind.

Evidence demonstrates that some primitive cultures had sophisticated knowledge about
the astronomical and solar phenomena. Stonehenge in England is a Celtic monument
that marks the solar solstices and the changes of the seasons. Likewise, the ancient meso-
American cultures were deeply connected to the Sun in their calendar as well as their

religions.

Meso-American Sun Worship

The early cultures of meso-America were perhaps the most elaborate Sun-oriented
cultures. The Mayans had a sophisticated society although much still remains unknown.
The supreme being was a sky god depicted as an old man. He also became the Sun god
and was believed to be married to the Moon. The Toltecs borrowed from the Mayans and
developed the myth that the Sun god died every evening and had to be resuscitated
every morning with human blood. Ancient mosaics show the offering of a human heart

to the Sun.

Aztecs drove the Toltecs out from their Mexican homeland but took on many of their
customs such as their calendar and their practice of sacrificing humans to the Sun.
However, the Aztecs took this sacrifice to new levels of morbidity. On occasion, sacrifices
of up to twenty thousand people would be performed. Tonatiu, the Aztec Sun god
pictured on the great stone calendar, was surrounded by fire serpents which defended
the Sun from his enemies at night. The battle between life and death, light and darkness,
was the entire foundation of the Aztec religion.

The Incas of Peru were much less bloodthirsty than the Aztecs, but they also had an
autocratic Sun god as a paternalistic deity. The Sim was the symbol of royal power and
the emperor was believed to be the son of the Sun. The Inca built their Sun temples so
that the sunrise fell on a golden disk which illuminated the shrine with a numinous light.

Chinese Legend

In China, there is a legend which tells of the plight of too many Suns and of the hero who
returns the world to balance.

A long time ago, there were nine Suns in space. Rivers dried gradually. Trees and plants
died as well. Everything was going to die. People did their best trying to save the things
in the world but could not. Just at that time a brave and kind young man came out
whose name was HOU YI. HOU YI wanted to save mankind and everything still alive no

matter how difficult it was and how big the sacrifice. He would give even his life.
Everyone was moved. Some people gave their ideas which would be helpful. Some went
back home to devote themselves to things they had just left behind and some youngsters
went ahead to join the activity.

HOU YI refused anything but food and water, brought his bow and arrows, and went
straight to the East where people believed Suns were bom and grew. He wanted to meet
the one who could manage the things related to the Sun, in order to ask him to cancel
some Suns so everything would be OK again. He went on and on, through many, many
lands, mountains, and dried up rivers; overcame lots of difficulties not even imaginable
today. At last, all he had was finished, no food, no water, nothing. He was exhausted.

When he was almost dead, he encouraged himself to stand up, stared up at the suns,
shouted to them "Why do you do things in this way? We don’t even touch you or disturb
you?" Then he laid down. He used his final energy to pull the bow, aimed an arrow at
one of the suns, and shot. One after the other he fired his arrows-0. Finally, eight suns
were shot down, only one was left. The universe restored its order. Everything became
alive. But HOU YI died without any regret. He had done his all for the whole universe
within which we still live .

Native American Legend

Arrow to the Sun—an Acoma Pueblo story.

A young woman in a pueblo is visited by a ray of Father Sun and bears a child, a young
boy. As the boy grows up he is ostracised by his playmates because he has no father. So
he goes to his mother and tells her he must find his father. He goes off and asks a farmer
who doesn t know, a potter who also doesn t know. Finally he comes to an arrow maker
who does know, and forms the boy into an arrow and shoots him on the long journey to
the Sun. °

The boy lands on the Sun but is told by his father, the Sun, that he must endure four trials
before he can be acknowledged as the son of the Sun. The trials are of endurance in kivas
of lions, snakes, bees, and finally lightning. With the last trial the boy is transformed and
can take his place alongside his father, filled by the power of the Sun.

The father and son rejoice but the Father tells the son that he must return to the Earth and
bring his spirit to the world of people. The Father makes the son into an arrow again and
shoots him off to Earth. When he returns he marries the Com Maiden and, with all the
pueblo, dances the Dance of Life.

Tapanese Sun Goddess

In Japan the Sun goddess, Amaterasu Omikami, is the centre of Shinto worship. She is
intended to bind the world together and maintain harmony among the gods, mankind,
and nature. The prominence of Amaterasu as the greatest reality visible in the heavens
symbolises the greatest reality known and revered on Earth.

An old Japanese myth about Amaterasu explains why the Sun is so important for life. It
also explains why many Japanese Shinto households have a rice-straw rope across the top
of their doors.

Many years ago Amaterasu, the goddess of the Sun, was abused relentlessly by her
brother and so she hid in a cave. In her absence, the world became consumed by
darkness. Other gods and goddesses knew that life would perish without the Sun so they

danced and played music to try and coax her from the cavern. Amaterasu was curious
when she heard the music playing. She proceeded to the entrance of the cave to see from
where the music was coming. When she came upon the musicians, a powerful god
pulled her from the cave while another god stretched a rope made of rice straw across the
entrance to prevent her from going back. The gods beseeched Amaterasu Omikami to
stay in the sky so that the world would remain light and never be consumed by darkness
again.

Ra: The Egyptian Sun God

Ra was the Egyptian Sun god during dynastic Egypt. The name "Ra" was thought to
mean "Creator" and took the form of a hawk or falcon-headed man. Ra travelled
through the waters of the sky during the day and through the underworld at night on a
barque or Egyptian river boat.

Some accounts of Ra's daily journey through the sky describe how he was born anew
each morning, grew through the stages of childhood, adulthood and old age only to die
at sunset. Other symbols associated with Ra are the scarab or dung beetle which
recreated itself by rolling its eggs in a ball of dung. The scarab was believed to roll the
solar disk across the sky.

Ra was believed to be the father and king of the gods. Tears fell from the eye of Ra.
These tears grew into humans and all living creatures. Ra presided during a golden age
period when men and gods lived together on Earth.

In Egyptian mythological structure, Ra was father of Shu and Tefnut, grandfather of Nut
and Geb, great-grandfather of Osiris, Set, Isis, and Nephthys and the great-great¬
grandfather of Horus.

Ancient Greece

Ancient Greece is perhaps the doorway between the human mythological relationship to
the Sun and a more logical one. According to Homer, Helios "rides in his chariot, shines
upon all men and deathless gods, and piercingly gazes with his eyes from his golden
helmet. He rests upon the highest point of heaven until he marvellously drives down
again from heaven to the Ocean." The image of the Sun in his chariot is seen over and
over again in Greek art and continues into Roman times.

The Sun in the Bible

In the Bible the Sun is an important symbol of God's illuminations as exemplified in
Genesis. "God made the two great lights, the greater light to rule the day, and the lesser
light to rule the night. . . And God set them in the firmament of the heavens to give light
upon the Earth, to rule over the day and over the night, and to separate the light from the
darkness" [Genesis 1:16-18].

4.1.2 Looking Back in Order to Move Forward

Why explore solar mythology in the context of a scientific project? One must remember
that in order "to understand where you are going, you must truly comprehend from
where you have come". Understanding what the Sun has meant to the human psyche
throughout the millennia is important for guiding scientific exploration into the future.
The exploration of the Sun will be as much a quest of mythological significance as it is an
objective scientific investigation into the Sun's physical properties.

62 • Ra: The Sun for Science and Humanity

Like the young Pueblo boy who seeks to know his father the Sun, the Ra solar project will
journey like an arrow to our Sun exploring the mysteries of its nature. There will be trials
to endure like the lions and snakes of technical challenges, economic difficulties, and
international co-operation. But in the end the mysteries that are revealed will be shared
with all peoples for the good of the world.

For this mission to succeed, we must draw on the mythic motivation that still drives our
quest for knowledge and adventure. For we are as much creatures of story and
mythology as were our ancient grandparents gathered around the camp-fire. Only now
the myths we live by are "economic development" and "scientific investigation" and our
camp-fires are computers and televisions. Consciously drawing on these mythic powers
can help motivate our generation to be "heroes" who provide good for all the people
through the exploration of space. Such psychic inspiration can propel this mission to
successfully realise our dreams of unravelling the mystery of our own star.

4.1.3 Heliobiology: The Influence of Solar Activity on Society

Not only has the Sun had important mythological significance, some philosophers have
investigated solar influences on social activity. In the 1920's a Russian philosopher
named Alexander V. Chizhevsky (1874-1964) began to develop theories about the
influence of solar activity on humans and their social behaviour. He belonged to the
Russian school of space philosophers and one of the main statements of this school is that
the Universe, Earth, and humans are constituents of one system which can be
characterised by life cycles and rhythms. He stated that "mass human behaviour is the
function of the Sun energy activity". Sun flow particles (or "z-flow particles," a name
given by Chizhevsky) have impact on the blood, nervous and hormone-endocrine
systems of different individuals.

Chizhevsky hypothesised that increases in the amount of the Sun flow particles within
peaks of Sun cycles caused an increase of excitability and aggressiveness of different
social groups on the Earth. The famous revolutions and wars of 1789, 1830, 1848, 1905
1917,1941 happened during the highest Sun activity, (period with the biggest number of
spots on the Sun's surface). During minimal Sun activity the social activity in society is
minimal, about 5% and during Sun maximums social activity achieves 60%. Sun particles
bombarding the Earth transform potential nervous energy of human groups into kinetic
energy that demands an outlet which results in revolutions and different mass
movements. According to Chizhevsky these social disasters change the velocity and
rhythm of the life period of different societies [Chizhevsky, 1937],

The ideas of Chizhevsky are under development now in Russia. His theory is being
applied for the prognosis of the further development of society, economy and
environment [Zhdanovich, 1994]. Special research has been made and correlation was
found between Sun activity and cardiovascular diseases [Atkov, 1996], Sun activity and
numbers of accidents and technological disasters, Kondrat'ev's economic cycles and Sun
cycles.

4.1.4 History of Solar Science and Observation

The history of more scientific observation starts with the Greeks who, six hundred years
B.C., made attempts to understand the Sun, the Universe, and their relationship to Earth
both through physical studies and philosophical ideas. The astronomer Aristarchus of
Samos measured the distance to the Sun through measuring the angle between the Sun,
the Moon and the Earth at a specific time. Though being underestimated to only 19 times
the distance to the Moon, a similar distance was adopted by Claudius Ptolemy of
Alexandria, and this distance was accepted for the next 1500 years.

Our View of the Sun • 63

In 450 B.C., Empedocles discovered that solar eclipses were caused by the Moon covering
the Sun, and in 350 B.C. Helicon actually predicted a solar eclipse for the first time.

In 1543, Nicolaus Copernicus proposed the Sun as the centre of the planetary system in
his famous book 'De Revolutionibus Orbium Coelestius', still using the underestimated
distance to the Sun from Aristarchus and Hipparchus. Only when Kepler stated his three
laws about the Solar System in the seventeenth century, did this underestimate give way
to a more correct idea. Kepler also stated that the planets do not have circular orbits
around the Sun, but elliptical orbits.

Sunspots were first referenced by Aristotle's pupil Theoprastus in the mid-fourth century
B.C., who also sighted the aurora. The first sunspot sighting happened in China in
165 B.C. As many as 157 records of sunspots seen by the naked eye were known and
scientists were well aware of their existence when the first telescope was discovered in
1608, allowing further and more accurate studies of the Sun. In Europe, records of
sunspot observations through the centuries seem to be lacking, due to the Aristotelian
view of the Solar System being strongly supported by the Church. Galileo observed
sunspots in 1610, using the telescope. Cristoph Scheiner, a Jesuit priest, observed the Sun
from 1611 to 1627, and both he and Galileo noticed that the paths of sunspots are not in
straight lines as the Sun rotates, but are curved, and they showed that sunspots are
confined to a band extending to latitudes of 30 degrees north and south. Heinrich Swabe
in 1843 announced the "eleven-year sunspot cycle" and also introduced the term sunspot
groups.

The aurora was given its full name, aurora borealis (northern lights), by the French
astronomer Pierre Gassendi in 1621, but many previous descriptions of the aurora exists.
Already in ancient Greece and Rome, as well as in early Chinese, Japanese and Korean
writings, auroral sightings are mentioned.

More detailed information about solar observation history can be found in Phillips [1992],
on which this chapter is based.

4.2 The Sun as a Star

In spite of the scientific means that have been developed since mankind first became
aware of the Sun's significance, the Sun is still full of mysteries.

It is amazing how little we actually know about our live-giving force. The standard
model of the Sun is threatened by the neutrino-problem, the origin of the magnetic field
is not well understood and the physics behind the eleven-years long sunspot-cycle
remains more or less unexplained. It is not clear on what time-scale and how much the
energy output of the Sun varies, the heating mechanism of the corona has not been
identified, and the physics of flares is a riddle to scientists.

In this section we introduce the Sun as it is seen by the scientists today. We begin with
how the Sun evolved and how it compares to other stars in section 4.2.1. Based on figure
4.1 we then explain the interior, photosphere, chromosphere and corona in the following
sections. Finally we describe solar activity in section 4.2.6.

m TKo Our* for ^ripnrp and Humanity

Corona

Fig. 4.1 The interior of the Sun [Beatty and Chaikin, 1990].

4.2.1 The Sun among Other Stars

In theoretical models of stellar structure and evolution, a star is taken to be a spherical
mass of gas (mostly hydrogen with some helium) compressed by its own gravity. Each
layer inside the star is squeezed by the weight of layers above it. The heat from
compression in the interior is transferred to the surface where it radiates into space.
Under these conditions of hydrostatic equilibrium, the radius of the star shrinks and its
interior heats up until thermonuclear reactions become possible at the centre. Initially,
the single most important nuclear reaction converts hydrogen into helium. Once nuclear
burning starts, the radiation becomes so intense that it can support the outer layers, and
the shrinkage slows considerably as long as there is fuel for nuclear reactions.

The luminosity of a star is proportional to the product of its surface and the energy
radiated per unit surface area. A star at a given temperature could be of any luminosity,
merely by being of the appropriate size. Nature however, does not make stars randomly
as was first demonstrated by Henry Norris Russell and Einar Hertzsprung in the 1920s,
described by the Hertzsprung-Russell (H-R) diagram [see figure 4.2],

If one accumulates information on the luminosity and temperature of as many stars as
possible, and represents each star by a dot in a graph of temperature (horizontal axis,
increasing to the left) versus luminosity (vertical axis, increasing upward), 90% of stars lie"
along a band called the main sequence in the H-R diagram. Hotter main sequence stars
are more luminous, and also larger, as one can see from the lines of constant size in the
diagram. The 10% of stars that are not on the main sequence mostly fall in the lower-left
corner of the diagram-a region of very high temperature but very low luminosity-and
thus of very small stars. These are the white dwarfs. A very small percentage of the total
fall in the upper right of the diagram, corresponding to low temperature but very high
luminosity-a circumstance which could only come about with very large stars--hence
their name "red giants". ' °

si cr

fYrir fl-i/-* C.im

As stars age, their luminosity and temperature change in a well-defined way. When the
luminosity and temperature of stars are plotted on a diagram, we see the points lying
along a path we call the main sequence. Eventually, stars exhaust their nuclear fuel and
shrink to become white dwarfs, neutron stars, or black holes, depending on their mass.

The Sun appears to have been active for 4.6 billion years which means it lies on the main
sequence [Noyes, 1982], about half-way along and has enough fuel to go on for another
five billion years or so [figure 4.3]. At the end of its life, the Sun will start to fuse helium
into heavier elements and begin to swell up as a red giant, ultimately growing so large
that it will swallow the Earth. After a billion years as a red giant, it will suddenly
collapse into a white dwarf-the final end product of a star like ours. It may take a trillion
years to cool off completely.

Qtottonortftttoor

ptanttaynabubt

100 000 r-

Fie. 4.2 The H-R diagram [Noyes, 1982]. Fig. 4.3 The path and position of our
& Sun [Noyes, 1982].

4.2.2 The Interior of the Sun

The Sun core can not be directly observed, as no radiation directly emerges. However, it
is possible to put together a picture of the Sun's interior with the use of the theoretical
solar core model. The theoretical model is a mathematical description of the way the
pressure and temperature vary with the distance from the core of the Sun to its surface.

The Sun's energy is released from the core by the fusion of four protons to form a helium
nucleus. At the centre of the Sun, where the temperature is calculated to be 15.6 million
Kelvin, the first stage of the nuclear fusion chain is the combination of two protons to a
deuteron. The second stage of the chain is the fusion of a deuteron with another proton
to form the nucleus of an isotope of helium, consisting of two protons and one neutron.
The final stage is the fusion of two such helium nuclei to form a nucleus of helium
consisting of two protons and two neutrons [Wentzel, 1989].

Most of the energy is produced in a comparatively small region near the Sun centre. Heat
is transferred by radiation in the deep interior to about two-thirds of the way out, then
convection becomes the dominant mode of transfer near the surface. The main zones of
the interior of the Sun are indicated schematically in figure 4.4.

Fig. 4.4

Principal zones in the solar
Bahcali, 1989) [Phillips, 1992J.

interior (based

on standard model of J. N.

are Sma11 packets ° f energy ' invisible a "d With no electric charge,
her they have mass or not is a question discussed by scientists. The word neutrino

C thTtZJ° m T neUtra ° ne u reIatin§ t0 the s P ecifics of tb ese small subatomic particles
hat are part of conserving the energy of the Sun. Their existence was postulated more

than half a century ago by Wolfgang Pauli, based on the fundamental principle of
conservation of energy within a system, and were first detected in the early 1950s

H V| U h n ° S P^ SS f 1 I ? Ug * 1 matter , anc * are not easily observable, and only in the 1970s
SunLsfon C reacHo V ns OP “P ab ' lili “ *° **** neutrinos emitted directly from the

Fusion reactions in the Sun can only be observed through the neutrino emission from the
main proton-proton chain reaction in the core. Thus, to obtain more information and
knowledge about these fusion reactions, and also to understand the before mentioned
energy conservation in the Sun, it is important to study the neutrinos and understand

their formation and existence. ers na

Fig. 4.5 Modes of oscillation in the Sun [Friedman, 1986].

Helioseismology is the study of solar oscillation. Modem helioseismology dates back to
1975 only when new technology and methods made it possible to further study the
spatial and temporal properties of the solar oscillations. This gives us the necessarv tools
to measure the depth of the solar convection zone, the internal rotation profile, the sound
speed throughout the Sun, the equation of state of partially ionised plasmas and the solar
ehurn abundance in the solar convection zone, by analysing the three different types of
sma 1-amplitude oscillations of the solar body about its equilibrium state:

• pressure-modes (p-modes), the pressure is the dominant restoring force,
the wave propagates by compression and rarefaction at the speed of sound
[Friedman, 1986]

. gravity-modes (g-modes), the gravity or buoyancy is the dominant
restoring force on a displaced mass of solar matter

• surface-modes (f-modes), nearly compressionless surface waves, also called
interface modes [American Association for the Advancement of Science,

1996].

Helioseismology uses all available pulsation data, including growth rates, p ases,
different modes-and not just observed frequencies-to search the internal structure and

evolution of the Sun.

In figure 4.5, contour plots of selected modes of oscillation of the Sun are shown Solid
lines represent expansion, dotted lines contraction. The longer the period of these
pulsation, the deeper within the Sun is the origin of the vibration. Though impressive
accomplishments have been made, there are problems related to background noise when
extracting information about the Sun's oscillation from measurements and observations
These limitations are however well understood [ESA, 1995]. As the science o
helioseismology improves, solar oscillations will give valuable information about the
interior of our Sun and about the processes happening within the Sun.

4.2.3 The Solar Photosphere

The photosphere is the first layer of the atmosphere of the Sun and the main part of the
visible and infrared light is coming from it. It has a very small depth of only 200 to 500

kilometres.

A typical granule, a convection cell in the photosphere, measures 110 km across, though it
is not clear whether there is a defined size scale for granules since they seem to be
steadily more numerous the smaller they are. The larger granules are bright, polygonal
areas separated by darker channels, called intergranular lanes. A typical distance
between two granules is about 1400 km. The smaller ones appear less regularly shaped.
It has been claimed that granules are on average smaller at sunspot maximum than at
minimum. The brightest part of a granule is generally about 30% brighter than the
intergranular lanes. This means the temperature in centre is about 400 K greater than in
the outer region of the granule. Their appearance is altered near sunspots, and they
become lengthened when they are in contact with the penumbral boundaries of spots.
Granule lifetimes average about 18 minutes, with the largest granules lasting the longest.
It seems likely that granules are rising convection cells of hotter gas and intergranular
lanes descending currents of cooler gas. There are strong horizontal flows from the
centres of granules towards the intergranular lanes.

Another convection is observed in the Doppler-shift of lines which indicates horizontal
flows occurring over tens of thousands of kilometres. These cell structures are about
30 000 km across and last a day, revealed by an outward, almost horizontal flow of
material from the centre of the cell to its sides, with velocities of 0.4 km/s. This
phenomenon is called supergranulation. The improved resolution of solar photographs
in recent years has resulted in the identification of a very fine bright structure in the
spectroheliograms taken in the light of weak Fraunhofer lines. This consist of tiny bright
points filigree, strung along the dark lanes between granules, frequently clustering to
form linear structures, called crinkles. The smallest elements are perhaps 150 km in size
and last for about 20 minutes. They are a few hundred Kelvin hotter than the
surrounding photosphere, and are associated with high magnetic fields. Connected with

68 • Ra: The Sun for Science and Humanity

these are faculae, the most conspicuous seen in the neighbourhood of sunspots. Others
occur at high latitudes and are therefore known as polar faculae. Both are associated with
high magnetic fields and both vary in number over the course of the solar cycle.

A comparison of the solar spectrum with the ideal case of a black body in thermal physics
shows a crude similarity with the radiation curve of a black body at about 6000 K This is
very roughly the temperature of the photosphere. Over the height range of the
photosphere, the temperature decreases from about 6400 K at the base to 4400 K at the
top. Beyond this level, temperature increases again, so that there is a temperature
minimum region. In visible light a point at the limb is at a level just beneath the
temperature minimum, so you see a less hot part of the atmosphere than at the Sun
centre, being less intense and somewhat redder. This decrease of solar intensity towards
t e limb is called limb darkening, it is very noticeable in whole-Sun photographs.

There are several possible line broadening mechanisms. The first is that resulting from
the motion of emitting atoms. The atoms move in all possible directions and any line will
have its profile broadened. This broadening is called thermal Doppler broadening. The
second mechanism is connected with the amount of time an atom spends in its upper
energy state. An atom making a transition from this state to the lower state emits a
photon with a small energy range. The spectral line formed is said to have natural
broadening. For certain lines, collision broadening is important. Charged particles do
not collide in a billiard-ball sense, but pass near enough to come under the influence of

t e electric field. The orbiting electron will gain a momentary perturbation. As these

collisions are random, the perturbations are random and so any emission line is
broadened.

The photospheric magnetic field is measured by the Zeeman splitting of certain
f n h f ° sph c en t c f 1Iy formed Fraunhofer lines. The largest field strengths occur in sunspots
(0.4 T). Fields exist elsewhere, and indeed it is likely that the entire solar surface is
pervaded by at least a very weak field.

A sunspot group generally appears on a magnetogram as a bipolar magnetic area, with
the leading spot having the largest field strength of one polarity and the following spots
slightly weaker fields of the opposite polarity. In addition to the active regions there are
many very small bipolar magnetic areas without spots. They even appear when the
sunspot cycle is near minimum and they have lifetimes of even less than a day. The
small-scale magnetic field is also associated with the filigree, which occurs where the
field is particularly strong (about 0.1 T). There are also small clumps of field
concentration distributed round the boundaries of supergranules. They are coincident
with structures observed in the chromosphere forming a chromospheric network.

4.2.4 The Solar Chromosphere

The photospheric Fraunhofer spectrum is, at the moment when a total solar eclipse
begins, suddenly replaced by an emission line or flash spectrum. The strongest emission
lines are the Ha line and the H and K lines produced by Ca* Therefore
spectroheliograms made in the light of these lines are used to study the chromosphere. In
addition to that, observations in the UV light can be made by spacecraft.

The outer edge of the chromosphere is very irregular. The edge is found to be made up
of numerous fine jet-like structures, the so called spicules. An individual spicule is
revealed to be a narrow column, a few hundred kilometres in diameter, ascending almost

. onmu 6 C iT° na With ve,ocities of 30 km/s. It is attaining an altitude of
about 9000 km and last approximately 15 minutes. Spicules are very numerous, but they

can only be seen in the solar limb. On the disk of the Sun there are small dark regions

(about 1000 km) visible, which are associated with an upward motion in the
chromosphere. Therefore these so called fine mottles are assumed to be the spicules seen
on the disk. They are located on the boundaries of the supergranulation of the
photosphere and have an average lifetime of 10 minutes.

An average lifetime of some hours and a size of 2000 to 8000 km have got the coarse
mottles. These dark areas form the "chromospheric network". The individual network
cells are about 30,000 km in diameter and last for some days if the chromosphere is quiet.
These patches in the vicinity of sunspots or other active regions are the most conspicuous
features of the spectroheliograms, particularly in the K-line images.

Sunspots, faculae and filaments are not directly connected to each other. They are
different responses to a perturbation of the magnetic field. The chromospheric features
and photospheric magnetic field are related on both small and large scales.

The most striking instance of solar activity is the solar flare, sudden release of energy
appearing as electromagnetic radiation over an extremely wide range and as mass,
particle, wave and shock-wave emittance. Flares invariably occur in active regions, being
most common and largest when the region is in a rapidly developing state. They can last
only for some minutes or for some days.

Although much information about the chromosphere can be obtained from images made
at the wavelength of lines in the visible spectrum, there is no indication of the connection
between the chromosphere and the overlying, much hotter corona. This connection can
be studied by observing the Sun in the ultraviolet part of the spectrum and in short
wavelength radio waves. The UV lines of the chromosphere, corona and transition
region tell us a great deal about the structure of the solar atmosphere.

Fie 4 6 The variation of temperature with height in the solar atmosphere
[Phillips, 1992].

From the base to the top of the photosphere, there is a decrease of temperature owing to a
decrease in the density of FT ions, reducing the ability of the photospheric gas to absorb
energy and maintain its temperature. However, as seen in figure 4.6, above 500 km
altitude the transport of non-radiative energy, whatever form it takes, leads to a rise of
temperature. This results in an increase in the ionisation of hydrogen, so there is a
greater number of free electrons and protons. The electrons are available for collisional

70 • Ra: The Sun for Science and Humanity

excitation, °f certain atom s and ions, which de-excite by emittme line radiation tw
em.ss.on lines include Ha and the H- and K-line of cT Ahho^h n

- *»> -

chromosphere ‘ Sm ' S assumed ,0 be responsible for the heating of the lower

° f ,h f ™*"«* field is though, to be important.

tTZdelSn^rWew HD ’ WV f°^

rensea elastic bands. The wave would give up some of its enero-v if *v, Q 6

^^Tbtg^gh^ and lhiS ««*V “ uM - a" e r

4.2.5 The Solar Corona

™*^rr,h:Xm°;'tre ea,,ng of the —* di “ a - ^ ~

^ b V ^orona.

only be observed during to,a, ec^pses. Th'e sXmofthtradfaL^up mhethtfo"

So^e^

=^rem d "

temperature“* 2 C ° nHnUUm ' ,he ”* ™ —ns te d^minoT

tOTs«s:«fi^SSsrf£-3

^•5^3SS=5is.*s=B

S^lSSiiSSSHS

te^mi nofckar where the slow solar wind originates from. It might arise
Te^een narrow o^n field channels between corona, loops, or "evaporate from iarge,

old loops.

The connection between magnetic fields and loop structure Jed ^ Propo^d
su^gested^duough^e^use of ^an^oref lEj^^ie^^b^o^the^issipation of Alvfen-waves

"• tsr^cssssi zg&zzpr Si

SESSSSrr,

times brighter in X-rays [Stem, 1996].

4.2.6 Solar Activity

1986L Sunspots appear darker than their surroundings because they are a few thousan

annulus around the umbra of a sunspot. Several sunspots can be seen in the full-d.sk
continuum image.

lZ?ol Z without

brighter^than'its'surroundings when obse J ed in the centre of a spectral line. An active
re Jon is essentially a collection of intense magnehc loops; they together from a ma^ie ic
u ?S1 or magnetic sphere of influence, in which the strong magnetism dominates the
U m ' f rViArtred particles in its vicinity. Energised material is also concentrated and

magnetic shape. They are the seat of change and unrest on the Sun. The interacting
magnetic ZL can, for example, trigger the catastrophicreease ofmgrtc «qy
stored within active regions, resulting in energetic eruptions, called solar flares. Indeed,

a " d radiation, as well as ,he eruptive

U-year solar cycle. Big active regions may gmwTo ^OW km in'dia' y ' he

dayT° m ° nthS ° r m ° re ' bUt Sma11 ° nes ma ^ a PP ear and disappear wiihTn TmaTtTrTf

Z '° abou h 5 T° km above the solar

strong spectral lines but not7\Z' 7 8 ' Li* 7 Can be observed in the centre of
thesectods appearbrightWhen selT”' s «=", b ^ d «>e limb of the Sun,
relatively dark and are called filaments Vl 8 /^ S ? ar disk ' the clouds appear
strong spectral lines, such as Ca K or H alDha^W 1 S 030 be SGen ° nly in the centres of
disk H-alpha image FUamPn^ .'nH P We Can See SeveraI fiIamer ^ in the full-

though some of them^disappear muchTs"^ remai " f ° r Up ‘° about two ™nths,

flares 6 The electr.caUy Sed e Ts^to make ' '° ap P ear aS a result of

above the Sun for weeks ormonto 1,1 h P 3 or lament can hover

gravity by the magnetic fitos to, 0*u ,upp P rted against *e downward pul] of
photosphere. X fonTtWn f lamenK 1 ,T bi P olar /t?'Ons in the underlying
magnetic neutral line centred between regioraof ODonsiT ° ma f netlc . loo P s ' alon 8 th e
the gas is held up by numerous magnette arches extend,ng [nTito*^ A PP arenl| y
top into a hammock-like shape. Despite its flamplilJl g ach sa 88 ln g at the

solar limb, a prominence is abn.u inVi flam f ,lke appearance when viewed at the

material. ThemameHcfields ,ha“ slolTT ^ ^ the s “rround,ng

and insulate it against hotter surrounding laW°r” °’ ’ amen ' also act as a shield

o” pafhcTe 5 3e°ST “?

Hundhausen, 1989, 1994 Lane 19951 SoW fl* u * k mas j motion [Feynman and

release o, energy Stored taX ma'nettfieldTthS toeTdto T ^ SUdda "
regions around sunspots. Usually solar flares arP nh • ar corona in actl ve

regions betoeen the u P pper chromSphte ““T*

from seconds up to hours. In the largest flares in 27 T „ bolar flares ma y la st

minutes. Such large flares only occur a be releaSed in a few

activity maximum. Many smaller flarec a ltbm a ^ ear or t™ 0 °f the solar

modem instruments at about 10 20 T These n d ° Wn t0 the HmitS ° f detectab iHty of
down to a few seconds their occurr^e TT 7T* ,aSt for shorter times

several tens offlares perda^ The L" f A *° f °u° WS ^ U ' year ^ P eakin § a t
unknown. Interesting features of the rp^ n ^ 6 1 1 aniSm of flare generation is still
formes shortly before the flare and d 8 ° mC u f a on 8 twisted X-ray structure, which
unsheared po't flare ^ 3 S7S,em bf

eCgTp^ To a CL f :; ld of f adaa ‘ Pb ”> d -i a “ 1 -"p-nha:

^connection of two shorter ones jus', tens of mini's

ffe^mlnTnd HmdhauKn^f cSSe'^' k r y Hnk With Solar acdvit y

and also the most energeHc events in rtL 2^f P c “^^tions of solar activity
unstable, carrying out bUHons of mns nf T SyStenf !‘ The ma § netic loo P s b ^ome
ouhvard-mov[ng g CME s ^^shetch the ^ 0 ^ m , “ lift off illto s P a «- The
bright rays rooted in the Sun. They can expand toTec " S T PS ' le ? vlng beh ind only

streaming outward past the planets and dwarfing every,tog tatoir part,""luh"
work only in one direction alwavc mrirrimo- „ °c in tneir P atb - bucb events

and never falling back in the reverse diction^MEs ari^v?s't n bubble rPl f ne | tary Spa “

magnetic fields expelled from the Sun into the heliosphere (figurl “ey o^nTxhibl

a three par, structure: a bright loop Smayteformed b"

above an eruptedl pro™^ Th' 1 ^ 1 "^ and P { ifts . of f ,i ke a huge umbrella in
rapidly expanding, bubble like sneu P F snowplough. When the

1000 kilometres per second. The energy of this mass mohon rs comparable to the ne,
radiated energy of a large solar flare.

^& m duted ,h A e

CSHKP model [figure 4.8; Lang, 1995].

Some flares detected in Yohkoh's soft X-ray images exhibit a helmet-shaped geometry
Some flare;5 detectea )i t edge or limb. In this X-ray structure, opposing

magnetic ^^^^^^^'llgg^^g^m^whic^the magnertc^tructureropen°and close
lends support to the CSHKP dare mooe 6 up-allowing the catastrophic

magnetic fields co S ejection, with its accompanying erupting

h "T!„tlv drives a CME outward. The corona may not be altogether surprised by this
apparently drives a CMtounv ejections occur in pre-existing coronal

s^mr,hltr4r:L b bri g rn hfr one ,0 several days before erupting. The he,me,
streamer is then blown away by the ejection and disappears.

, . , i ;f rup Q rnuld be magnetically controlled and driven as the

Therefore, it looks as whUe wg belie v e mag netism to be the ultimate source of

theoretical mode s gg • ' CME Qne has ever measured the predicted

energy involved in both solar flares ana ^ , eruptive outbursts on the Sun.

depletion of magnetic energy that su PP ose ^y ^ be all the magnetic action

Perhaps the inshuments are <he

occurs in the unseen coron . ^ litt , overall change in the magnetism would

-r" if'tWsIs^heca^ why aren't the erupflons more powerful and why don't
they occur more frequently?

Fig. 4.7

Coronal mass ejections [Lang, 1995],

Ho"e^rJZ m J ' r Ur " re§i ° nS ° f Str ° n S ma 8-tic shear in the photosphere
owever many sheared regions never erupt, so contorted magnetism seems to ho a
necessary but not sufficient condition for solar eruption.

Debates continue to rage over exactly what strikes the match that ignites explosions on
the Sun. Magnetic fields coiled up in the interior could bob into the^corona and^ nZ.rt
W.th pre-existing ones, or existing coronal loops may be brought into contact bv X
shearing and twishng motion of their photospheric footprints The enmHrm u/

fnsiSthe'sun 5 ' ft co^beT" 1 * C ° r ° na ' l0 ° PS Whe " lhe y can “> rrturnTack

but no onetows for sure ' mteraCKon 0r Emergence of coronal loops,

e Sun of pent-up frustration and to relieve it of twisted, contorted magnetism^ ? §

Billion-ton bubbles of hot gas grow larger than the Sun in just a few hours This time
equence of coronagraph images, covering just over 4 hours, illustrates some of the
principal features of many CMEs the presence of a hHo-ht ™ . i r

followed by a dark cavity, under which is visible a bright foop-like 6 ^^^^ identT^'
with erupting prominence material. According to a version of the CSHKP mini
proposed by Peter Shu-rock in 1968, the magnetic reconnection results fn high-energv
particle acceleration of two bright ribbons in the chromosphere. §h 8y

Fig. 4.8 Composite eruption model [Lang, 1995].

In this cross-sectional view, a CME, and the erupting prominence that follows it blast an
into previously dosed magnetic fields,

the reconnection site give rise to intense, impulsive radio and hard X-ray radiation.
When the reconnected magnetism regroups to form closed structures, post-flare loops
shine at soft X-ray and H-alpha wavelengths.

4.3 Interplanetary Space

This section introduces in the basic physics that is known about the solar wind, the
interplanetary magnetic field and several dynamic features like magnetic clouds and the
pro^gahW CMEs. Furthermore a brief description of galactic cosmic rays is given.

4.3.1 Solar Wind and Interplanetary Magnetic Field

Solar wind is an invisible flow of superheated, charged coronal gas flowing continuously
out of the Sun. The particles traverse outwards into interplanetary space, eventua y
hittine bodies. Solar wind consists not only of particles from the Sun, that is plasma
consisting of protons and electrons, but also of particles from the interplanetary medium,
including comets, asteroids, and atmospheres of planets and satellites.

Solar wind composition can be determined in great detail by observations^ If the
composition of the corona-where the solar wind originates-is known, assumptions can
be made also about the composition of the interplanetary medium. Generally, solar wind
consists of 95% protons (FT), 4% alpha particles (He~) and 1 /o minor ions: carbon,
nitrogen oxvgen ? neon, magnesium, silicon and iron. The energy of the ions range
between 0.5 and 2.0 keV/nucleon, at a density of 1 to 10 particles per cubic centimetre.

There are two types of solar wind, slow and fast. These are affected by the solar magnetic
Sid and as thLlow and the fast solar wind leaves the Sun surface, the two interact
because of the rotation of the Sun and create compression and rarefaction, forming he so
called corotating interaction regions. The fast plasma in the stream overtakes the slower
plasma and collides with it. The plasma and the magnetic field are compressed, the
plasma is heated up and pressure waves are produced. These pressure waves enhance

othe^° f ,he faslplas ™ behi " d

in the surrounding plasm. S" . ^ P ' a ? ma ,s hl 8 her than the *P«<1 of sound
enhanced pressure den^tv ?h ° n * “J* 6 . of ,he pressure wa '' es a front shock with

back shock with dKteas J proiSnem^d nltT d 'fT’? ° n the back ed S e a
1986, p.191], P ' d ty and P^sma velocity is created [Marsden,

IhhoughTset^^^ k ™ /s -

fieM > outward^form^g I whanscad!«l the ^nterphmefa^magnetic

the solar wind moves out almost radially from the Sun the rotation of ^ c ‘ h “£ h

Fig. 4.9

PII 7 v!L.? froz . en into a radial solar wi
speed of 400 km/s. [Kivelson and Russel, 1995]

expansion at an average

— £E£S ° f SO ' ar Wi " d P ‘ aama ' * ™F -rioVat an

magnettsed body; a comet, an unmagnetised body with negligible giaJhv ' 3
unmagnetised body with an atmosphere (e.g., Venus}. g S ^ S ty ' and an

4.3.2 High Energy Particles and CMEs

Very high energetic electrons and protons T^pa^

zs, ssMSiffiSSK-*-- 1 *

encounter the Earth in twelve to twenty minutes.

Fig. 4.10 Interaction with bodies in the Solar System [Biemat et al, 1994],

ThP field lines the energetic particles follow must originate on the visible side of the Sum
The held lines tne energe F move across the field lines in the mhomogenous

Origin of CMEs at the Sun* fromThe corona "ofthe Sun to theW.
Thev'carfbe'taken^bubbles in *e solar wind. These bubbles are reservoirs of plasma
«a P W by closed field

formation moves i wrth, velocmesupto ««*»« are Vacterised by stiongly
enhanced helium abundance and bi-directional streaming of supra-thermal elections and
energetic particles.

4.4.3] , /tp^ cVtnrW<; which C3.U.S6 sudden storms on the Esxth.

^—•— wi, s

tp disturbances and with radiation hazards at Earth. The geomagnetic index use
^quantify the magnitude of geomagnetic storms [section 4.4.3] is highly correlated wit

ro anH Hnmanitv

A magnetic doud " *«-

approximately 0.25 AU at 1 AU and a low pIp h- 100 ma S netlc field direction over
Clouds have a loop-like structure and i° , f C *T and P rot ™ temperature. Magnetic
1990] as illustrated m * the Sun f BurIa S a a " d Lepping,

constant speed which depends on the surrounding^ ^ e *P ec * ed to P ro P a gate with
1996], but the direction of [Umar and Rus t

direction [Smith et al, 1996, Vandfs et al 1995 199^1 ^ ™ fr ° m the radial

for solar-terrestrial studies first because of thoi' * r Magn ^ hc clouds are ideal objects
Secondly, their extended o( ,!SSri;“ d their longevity of passage,

combined to smooth the variation of the field (»F n a,ad n ° rthwar d magnetic fields
to the magnetosphere very Vel ° ci ^ make « effects

1994], y uwcompare d to its ch^ctenstic Hme-scale [Biernat et al,

Q EARTH

Fig. 4.11

A schematic showin
flux rope [Biemat, H.

g a magnetic cloud modelled

K. et al., 1994],

as a toroidal magnetic

4.3.3

Galactic Cosmic Rays

Galactic cosmic rays (GCR) are extremely energetic f~10 9 pV^ ^ ,

mostly of protons, that enter the heliosphere torn the intemteUar medium" ' C ° nSiStmg

In the measurements made on Earth and t

intensity of the GCRs in the Solar System is reeulahS?^ * ^ been found that the
the maximum intensity of GCR occurs during the^nimum 0 ** by S ° lar activit V :
mirumum mtensity during the maximum solar activity actlvit >'' and th ?

form. As the GC , R in their un changed

magnetic field lines in the solar wind the bes location to? consider ably by the

is near the poles of the Sun, where the solar ma™ h relatlvel y ^changed GCR

equator. ' the Solar ma g net,c held is not as strong as near the

4.4 The Sun-Earth Interactions

interaction in between. Furthermore their dynamics
magnetospheric storms and auroral lightning.

will be described including

4.4.1 The Earth as a Magnet

For the space physicist, the was the first to

nn^^K'^n« «s of^restria, ma g netism had been
utilised much earlier by the Chinese in primitive compasses.

in firs, order the malefic field P

north-west of Hudson Bay. THe north g P P ^ t of Little America. The

” <* - *“* in th * — hemisphere '

and away from it in the southern hemisphere.

The magnetic field can toUhquW and »

are the source for the mam fieid. 8 . g d ^ e tQ curren ts in the ionosphere. The

changes are called secular chan 2_ tQ the secu i ar changes) but are also very weak,

variations there are very fas ( P the transmission of radio and television

4.4.2 The Magnetosphere

The area around the Earth governed by the E gJ^ t S a^finiHcms of some of the most
magnetosphere, and to boundary the magnetopaa^ Short drfirutio^ ^ ^ ^ ^

•» •*“ hom Kivelson and Russel 119951 and

the Space Physics Group of Oulu [1996].

The existence of the magnetosphere is of £

planet from the high energy P* r 1 caTfies al the IMF [see section 4.3.1], modify the
solar wind and the magnetic hing^t in the dayside and creating a long tail

form of the ma ^ et ° SP uf Vp^A s a^onsequencf, the distance of the magnetopause from
(magnetotail) in the nigh si • ^m) j n the dayside, while the tail is about

the Earth is only about 10 Earth radu (Re Tr^LiL a, more than 1000 Re). In

10 times longer (it was registered y boundary called the bow shock, is formed

front of the dayside magnetopause: another boundary^called^ ^ ^ ^ ^ ^ x

speed of sound waves in ,he solar wind plasma ,S 60

km/s.

The magnetosphere is fil f le £^the plasma in the
solar wind. Because of the m g P sca i e sunward motion called the

closed tail field fines is forced ^ow the solar wind drives this convection is

magnetospheric convechom The e usuaU y assumed , h a, the Earth's magnetosphere is

c _ c r] T-Tnmanitv

SOU lhw J^^

and a so “

Fig. 4.12

«* Won reconnection

sprfinn l n; c • \ magnetic storm and occur several times therein fsee

-eutral sHee,\cunent S) are^poTSylS^dthe^ ““ be '° W

thinning during the substorm growth phase. ' P sheel 15 lypicall >'

^S3SSSH:

The ring current flows around the Earth in a circle at distances of about 4 to 6 Earth radii

HHHHillP

no clear distinction between them Thp n l«m a A a •. • j ”, e because mere is

significantly higher than outside, because field lines at Wg^Hmde's a^contS to

Our View of the Sun • 81

the magnetopause and are thus open
ionospheric-supplied plasma is lost).

to the interplanetary medium (where the

Fig. 4.13

The Earth radiation belts. The top panel shows the of

The trapping regions of high-energy charged particles surrounding the Earth are called
b l^een 1 "lTnd^S 3 Rein thetquaLrS with energies

stable population but t is sub^t o occas p ta tHs ion is the decay

section 4.5.

This was only a brief introduction to the magnetosphere. For more informatxon and
further reading see Kivelson and Russel [1995].

4.4.3 Magnetic Storms

Geomagnetic storms are initiated when enhanced energy is transferred from thei solar
wind/IMF into the magnetosphere, via magnetic field merging Isee sec on . . ,
leads tointensification of the ring current. The ring current, can be measured with the

K e q t X ,^r^ 4 \amu e dS' £ d V S ,i ° bt r d fr ° m “»T stations near

ris=fi~||||sgsH

definition has been proposed by Gonzales : 7 Y * ' Th followin g storm

»SSsHl£iHfF“f h ~~ assas

fsAL o sr«ar key ,hres,Jd of ' he '.I™ ir bT,' iS g

major n erta^mS fheToteS v*TC P “ ia “ y ' he largeSt °" eS ' ° f,en ‘’“S 1 " with
referred ,o SO “ ,MF “

become greatly disturbed broadening ar*A a uring a storm, auroral ovals

on the nightside TWsTrin^theT* .T™?"' 8 ‘“wards the equator, particularly
section 4.4 5) aUr0ra ° f middle and low latitudes 1=““

Storms are typically divided into three distinct phases according to the signatures in Dst:
Initial phase

Lasts from minutes to hours. Dst increases to positive values up to tens of nT.

Dayside magnetopause is compressed inward (perhaps by several Re).

Main phase

' hundredTof h nT f ^ h ° Ur '° SeVera ' h ° UrS ' DSt Can rea “ h "“S a «ve values of

• Ring current is built up by multiple intense substorms.

Recovery phasp

• Lasts from tens of hours to a week. Dst gradually returns to the normal level

• Ring current ions are gradually lost.

Geomagnetic activity as a whole has a seasonal variability with maxima at th. •

This is especially true for intpn^ cfnrmc i: . Un maxima at th e equinoxes.

within the solar 7 cycle on^so^ mtenSe Storms sh ™ *w° Peaks

described infection 4.5 f PfaCtlCal im P ortance produced by the magnetic storms are

4.4.4 The Ionosphere

tdt^lhnu'lt^s'cornposedo'fitnvlsed^iTfplasn^a)^lasm^co^T 1 * 6 3 ' •I'T" at ° Ve ab ° U '
but in ionospheric phenomena charged panicles have fhe main ChTrg^cZ

are affected bv electric and magnetic fields. They also carry electric currents and hence
cause magnetic fields themselves. Effects of these fields are discussed in section .

The origin of * the; fo^phSids

“?fom eSgSc r" atmosphere is a filter which stops all short wavelength

radiation from coming to the surface of the Earth.

rSSHSSssrSS

[Akasofu and Kamide, 1987].

fonosphere [see figure 4.14]. Maximum of electron density vanes typtcally betw
and 300 km, but below that there often appear several bumps .

light.

Fi * 414 atassfisssaswass ssssswass

19871.

The ionosphere and magnetosphere are not separate parts of the Earth's near space. They
lod with *»ach other bv electric currents which transfer energy between them.
Currents exist at all times but during magnetic storms and auroral substorms they are
stronelv intensified. These currents are the result of particle streams which lose

rJL^due to collisions with atmospheric particles. Collisions cause energy transfer
frorrfthe current to upper atmosphere and ionosphere due to heating at heights aroun

and is then released as elfctrom^etirrldfation^Th" ^ ^ molecuIes '

pare,cle reservoirs in rhe SST"? T ^ ° n char ® ed

ionosphere" ““ "* ™ S is * he "gton of ft

currents (MC^BiAd^ct«t^ttraH and fh i0n >f Phere are Ca " ed field aligned
Hnes. Downward and upwaTS/d 0 ^ ^ m^Md

ionosphere by Pedersen currents flowing pfrallel with Ap ^ ^? nnec ; ed in ^e lower
meridian. They are closed in the mag^o^re rn ^ f * 5Urf ** e md ma S netic
electric field is perpendicular to the maenetk- fi^lH °™ e J part of this loop the

particles in east-west direction (Hall currents) Thpc Causin S rdt currents carrying
oval are called auroral electrojets [Meng et al 1991] ^ ^ ^ the aurora '

The equatorial electro je t

current turns to the southland it L comecTed^ 6Venin * §ide tlais
northern side of equator. Currents in the Jrh <V field lines to current system in
[Akasofu and Kamide, 1987], gHt Slde are sma11 due to lack of charges

4.4.5 The Aurora

around the Earth. P & changes m electric and magnetic fields

During magnetic storms energy originating u, i

magnetic field, is stored in thf magnetic lid 0 T th.% S 1" Wi J d and interplanetary
atmosphere carried by accelerated charged particles fm l^i 3nd then reIeased to the
quief condiHons electrons don't have typicahv hieh y e ' ectrons) ' D "ring so called

■n magnetic storms they can gain ve^y high enlrmes The Ho ’T: er 1 wh ‘ ! " accelerated

^:z h z?:zz j,ed s,ate reiease ^oma^Ldiir^t,^

Auroral luminosiVacSy^oiSt^ofnumeroS lllustratl "K the sha pe of current sheets,
bands. Green 5577 nm line is SoMham.n hiXr h Thf d emiSSi ° n lines and e ™ssion

concentration" '° WeSt aWhldeS Iines b -me do^inanTZ' mZmtin^"

and ionosphere. ThTs'lntensMct Hon^an obsZId 5 - 5 " h betWeen the m ‘’S'ietos P here
ground. Luminous effects show currents alone th^ magnetic field °" *e

phenomena are polar auroras. They are caused hv n .fV rora oval - Other auroral
solar wind along field lines reconnected to the [Menf at“ ,m? 8 d ‘ reCt ' y fr ° m the

Heating Lu^y^Ztan^^ Zthe"ionosphere ^° S P here and hum an activihes.
hence winds. Currents also cause momentum exchan J? < tem P eratu re gradients and
particles resulting in the same effect. One Carriers to neutral

following current from dayside of the Earth tn P f , f , that 1S neutral particle flow
magnetic fields which can^Luseeffects^ T' AUf ° ral CUrrents cau ^
section 4.5. on the S round and near space as described in

4.5 Effects of the Sun on Earth, Humans and Technology

The Sun is the primary source '" er ® y a m au u ofMlm activity, it is reasonable to
functioning. As solar irradiancech g possible climate change could imply,

think that these variations can affeC ^^ C ironme P nt al disturbances such as the shift of

«—" 1

related social impact.

Radiation effects on humans On Earth, tola^ activity^can

fatal cancer increases as doses and to ^ alsQ more direc t and less controversial

cause many serious medical proble _Th ^ magnetospher e's protective barrier and

whe^solar^Mes^ o^urs The ^ un and^orr^^'cation inte^erence

“e 1 eSTre therefore extremely important to take

into account in space missions.

These are the main reasons why we^should d Siolo^ by obtaining

new Kom spaTe'and solar probes missions.

4 51 Effects of the Sun on Earth's Climate and Biosphere

Since the Sun provides the drtrt^S “ preTJl

output are obviously a potential m betwein the Earth's temperature and

there is already statistical proofs of a is thoughl lo be mainly influenced

variations in the solar cycle. H °7™'we need to know the

by atmospheric concenhahons <rf t0 Hs change in order to analyse

on the variation of Earth's climate.

Environmental and social impltotiOTS M a tMu^chmate dmp ba^

starvation and loss of biodiversity.

ninhal Climate Change

Vjiuuai - « J —

It is thought that real wanmng of the globe of 0.3 C to 0^6 c^has^te^^p ^ answer , he

from hundreds of millions of years to a .>! 'elacial cycles when the climate was

have been the 100,000 year ^“““"nd^lmbrie 19791- Global surface temperatures have
mostly cooler than at presentjlmbne and lm with large changes

?s srea * as io ' 15 ° c in some

^Sie^Xbtbmde^ons of the^or^m Hemisphere.

In an unperturbated state the sola , r “ d here'by^^utgJing ^**100“^^
atmosphere is balanced at the “P”''” ra *ationis reflected back to space. Of the

but most is absotbed by ,he land ' ocean '

Sttaiisii

s~a=3sa=s=s

Effects of the Sun on Earth's Climate

SS^IteX^ £* l0n8el “"T" 1 * usi "S indicators have been

change in the future will be ve^ dimate

believed to MtiateV^XV t ^ ,he Ear ' h ' S orWtal Pieters are
100.0M years These ^ 10 '° 0 ° '°

Changes" tn^otal solaHrSce^r'Ihe Iou“t“ ® ^ to

SSssa 2 a ^ a " s y - ave ;r d fo r*

Sunspot cycle (cZ'e Change lM^ ' “ ' hen ‘ rrad,anCe has increased *« '° *e

Correlatio n between Climate and Solar Activity

^^^ia^nce^fiave^been'pu’e^ctsely^ea^ured^f^ n ^ e ^ ra * e< ^ T* r6S ° Ived solar

instruments. On the other hand the ZArichTh 3 nU [ nber of dlffer ent space-borne
number back to 1700- epochs of mavima • • serva ® r y reconstructed the Sunspot

estimated and tab.es o/s/dt informaHon ^7^ by^e “““ ^

^octaHo^e^'^?ro^te7 P vSon7 0r , d h a 7 SO ' ar aC,iV “ y indica,es a S° od

record, although S2,2™less^“oWous ffl£T ^ t** ^ ^
than for the modern instrumental record In 1991 Lassen andR pre , M " menl ‘ lj P eriod
this relationship correlating the tetWtum devL^Z 1 M Fn '?- Ch ™ tensen sho »ed

Reid 1991, Tinsley and Hee.is 19931.

t?:Z “ S ' ^oba. average sea surface
temperature, etc.).

i t; j il 0 ticpfulnpss of solar cvcle correlation studies, noting

observed, but unrelated to solar forcing [Dunkerston and Baldwin, 199 ].

new data from Earth, space, and solar probe observations.

Implications of Climate C hange to the Biosphere

a new climate distribution implies the redistribution of biomes (terrestrial regions
A 't7 t „dZ^ certin types of life) associated with loss of biodiversity, the change in

cause the rise of sea level.

, __ in flip Farth's history have been associated with shifts in the

]e^raphk7istribuUon of terrestrial biota. For example « :h *.

* r. „.-ii nr , r th of the current timber line during Medieval Warm Epoch (»UU to
l“.)7a time when temperature in that region was about 1 «C warmer than today.

A shift in the geographic distribution of biomes is a long-term (decades to centuries)
response to climate change and

Itat^K^Xac^primary confers of the biogeochemistry o,
ecosystems.

Photosvnthesis and respiration have different optimum ranges for temperature and

optimum range which is s P e< ^ ? di ^ a te changes to individual plants is

selectively favour^ corrununi ties; complex interactions of ecosystems

mustreadiust to new conditions as a result of changes in competitiveness of species. The
. .. n hvsical change the stronger the ecosystem is affected. However, the mos

mmolex ecosystems such^s tropical forest and coral reefs, are well adapted to constant
weather conditions; little changes in the climate could dramatically impact these fragi e
ecosystems with consequences of loss of biodiversity.

On the other hand, in order to predict climate variations, the effects of terrestrial
ecosystems changes on the climate change must be taken into account. Some induced
changes of ecosystem structure and function are expected to feed back to the climate
system. For instance, the warming of high latitude wetlands will almost certainly

increase the production of CH 4 and as it is released into the atmosphere it will accelerate
warming.

One of the more generally accepted conclusions of the general circulation climate models
is that as average global temperatures increase, the hydrologic cycle will speed up,
increasing global precipitation. As temperature and precipitation patterns change, so will
soil moisture and the timing and magnitude of runoff, with possibly adverse effects for
many of the world's important agricultural areas. One likely consequence of these
changes would be that the demand of water, especially for irrigation, would increase in
some regions. As pointed out in the last part, the combination of temperature, moisture
and water supply optimise plant production. Therefore, these variables will drive the
new distribution of agricultural production, how crop yield will change, and also forestry
resources. J

It is highly likely that the global-mean sea level has been rising over the last 100 years.
The estimates of different studies ranges from about 0.5 mm/yr to 3.0 mm/yr. There are
two major climate-related factors that could possibly explain the rise in global mean sea level
on the 100-year time scale: (1) The thermal expansions of the oceans. Density is inversely
related to temperature, thus, as the oceans warm, density decreases and the oceans
expand and the sea level rises. (2) A possible increase of global temperature will cause a

direct effect on retreating glaciers, small ice caps and polar ice sheets which will cause the
rising of sea level.

Based on the record of the past, there is a little doubt that global warming will result in
different distributions of marine planktonic organisms than those of today. Changes in
temperature and precipitation will have an influence on the circulation of surface waters
and on mixing of deep waters with surface matter. Changes in circulation and/or a
restriction of the mixing could reduce ocean productivity. As in terrestrial ecosystems a
global warming will redistribute production as a consequence of different spatial patterns
of physical conditions. Since fish concentrate in rich plankton production areas, fishery
activities would have to change their common areas of activity with possible
consequences of social and state conflicts.

The adaptation of our society to these changes will depend on the degree, the sign of

regional change, and the capacity of the particular culture, that is the technological
development. °

4.5.2 The Effects of the Sun on Humans

The Sun affects both people living on Earth and astronauts in space. These effects will be
discussed below.

4.5.2.1 The Sun's Effects on Astronauts

The issue of radiation may be the "big show stopper" in respect to long duration manned
space flight. The trapping of ionised particles by the Earth's magnetic field in the Van
Allen belts provides a shield against deep space radiation. Such ionising radiation exists
in many forms-high energy protons, heavy ions, and electrons-and may originate from
solar flare (solar energetic particles), the particles trapped in the Van Allen belts and
galactic cosmic radiation.

The effects of this deep space radiation on the human body are not well known because
all the past human space flights, with the exception of certain Apollo missions, have been
in LEO, which is well below the van Allen belts (except for the South Atlantic Anomaly).
The Apollo missions minimised the dangers involved with radiation, by avoiding periods
of solar flares. Some scientists believe it is unethical to send humans beyond LEO, as the
consequences will range from an unacceptable increase in tumours to possible death. It is
not known what type, if any, of shielding will successfully protect humans in this
environment. Ironically, the more shielding you use the greater the danger from
"secondary" radiation becomes. Impinging particles impart their energy to molecules in
the shielding material, rendering them, in turn, ionised.

Exposure to space radiation is painless. On a long duration mission to Mars, cosmic-ray
particles will pass through every cell in the body; however no immediate ill-effects
among the crew are likely. The risk of getting cancer in the years to follow, increases.
Radiation effects on humans are generally placed in two categories:

1. Acute, early effects of radiation exposure occur within a few days or less.
These are usually associated with exposure to a high dose of radiation over
a short period. Indicated by symptoms of radiation sickness.

2. Delayed, late effects may occur many years after prolonged exposure to
radiation at a low dose rate. These effects include cancer of the lung,
breast, digestive system and leukaemia.

Doses in the range of 100 rem to 200 rem (rem is a common unit of dose equivalent, 1 rem
= 1 rad = 100 ergs/gram = 0.01 Si) generally cause nausea and vomiting within a few
hours, which may be accompanied by discomfort, loss of appetite and fatigue [Churchill].
These symptoms disappear after a day or two, but may recur after a latent period of
about two weeks. There is little chance of death from exposure at this level.

Doses in the range of 200 rem to 1000 rem are very serious and require medical attention.
The initial response to radiation in this range is similar to radiation at a lower dose
exposure, and diarrhoea may occur. After a latent period of two weeks other symptoms
may occur including haemorrhaging and hair loss. The dose has caused serious damage
to the blood-forming organs, limiting the body's ability to fight infection. Doses above
600 rem are generally lethal, but recovery is possible with adequate medical care.

In space, doses of 1000 rem are possible in cases of large solar mass ejections. Provisions
for a "storm shelter" or other safe havens are essential for extended missions in space.

An astronaut's chance of fatal cancer is increased approximately 2% to 5% for each 50 rem
exposure during his/her career. In concrete terms if 100 Space Station astronauts are
exposed to 100 rem during a one year career in space, then between 4 and 10 of those
astronauts would be expected to die of cancer resulting from that occupational exposure.

4.5.2.2 The Sun's Effects on Humans Living on Earth

The Sun can have many negative effects on humans. Most commonly known are the fact
that looking straight into the Sun can cause blindness and that UV radiation causes skin
cancer. There are also a number of medical effects for which the correlation with solar
events can not be explained. Effects like these are studied by a branch of science called
biometeorology. Examples of these effects include:

90 • Ra: The Sun for Science and Humanity

• Sudden, unexpected death in epileptics following sudden intense increase
in geomagnetic activity [Pycha et al., 1992]

• A drop in human immunoglobulin levels at the end of the 11-year sunspot
cycle [Tisdale, 1995]

• A rise in intraocular pressure in healthy people during periods of increased
geomagnetic activity [Tisdale, 1995]

• Correlation between increased solar activity and heart attacks, epileptic
seizures and growth in hormone levels

There is a correlation between periods of geomagnetic storms and an increased number
of angina heart attacks in patients with high blood pressure [Atkov, 1996]. Geomagnetic
storms occur on average once every two months and are the result of solar activity. The
connection between angina heart attacks and geomagnetic storms was discovered while
trying to determine a correlation between medical conditions and weather patterns. It
was found that geomagnetic fluctuations can cause heart attacks in certain high risk
groups, such as elderly patients with high blood pressure. The full extent of this
relationship is not well understood, but it has been discovered that angina attacks are
most likely while entering or leaving periods of geomagnetic storms.

If sufficient warning of such storms could be given then doctors could prepare their
patients who are most at risk, by giving them the appropriate drugs. If an early warning
system like this made the information available to the medical community in real time,
then deaths resulting from angina attacks would be reduced.

It has also been shown that UV light from the Sun can activate the human
immunodeficiency virus (HIV) [Sun Exposure and HIV Activation web page]. These
findings were the result of tests on laboratory mice which were introduced to the HIV
virus, and subjected to UVA and UVB. While awaiting results of further test it was
recommended that people with HIV should avoid excessive exposure to sunlight and
wear a SPF 15 or higher Sun block.

4.5.3 Technology

At first glance, the Sun s effects on technology do not seem too obvious or too severe.
However, the Sun's influence on the space environment can present tremendous hazards
to spacecraft. Earth-bound instrumentation and communications in space and on Earth as
well.

4.5.3.1 Effects on Spacecraft

Great pains are taken by engineers to overcome the changes that the Sun effects on the
space environment. Even so, the Sun can cause problems that degrade or even
prematurely end a spacecraft's lifetime.

Atmospheric Drag

Solar emitted X-rays, extreme ultraviolet radiation and charged particles that intersect the
Earth, deposit their energy in our upper atmosphere. During intense geomagnetic
storming or periods of increased solar activity, this deposited energy forces the
atmosphere to heat up and rise. Satellites and orbital debris orbiting through this heated
atmosphere experience varying atmospheric densities which result in a loss of orbital
altitude along with pointing perturbations. This atmospheric drag will make the object's
position somewhat lower and ahead of where it was expected to be. These effects may
even cause early and unplanned re-entry of orbiting objects into the Earth's atmosphere.

Our View of the Sun • 91

just as Skylab did in 1979 [Worden, 1996]. Atmospheric drag will delay acquisition of
LEO satellites, expending valuable antenna contact time. It also can necessitate
additional manoeuvres to raise the altitude of the spacecraft before atmospheric re-entry.
Atmospheric drag also complicates orbit debris tracking necessary for collision avoidance
missions. Since an estimated 25,000 pieces of orbital debris are created in Earth orbit
monthly [Wilson, 1995, p. 158], collision avoidance is more and more important for new
payloads and piloted missions.

Surface Charging

Low-energy electrons deposit their charges on the spacecraft surfaces and over time,
these charges build up. Eventually they will produce a discharge that can cause
erroneous signals to be read by sensors and can permanently damage electronic
components and photovoltaic cells. These effects are observed to prevail in high
equatorial orbits along with low polar orbits [Lemke and Mendell, 1996]. More
information on surface charging for interplanetary missions can be found in section 6.1.4.

Single Event Upsets

Heavy ions and high energy protons emitted from large solar flares occasionally will
impact spacecraft. These particles have sufficient energy to actually pass through the
spacecraft's structure and change the spacecraft s chemical bonds [Lemke and Mendell,
1996]. If these particles happen to come into contact with sensitive electronic
components, single event upsets (SEU) may be experienced. An SEU can re-write on¬
board computer memory by replacing l's and 0's or may actually cause erroneous
commands to be executed by the vehicle with unpredictable and perhaps catastrophic
effects. An SEU is suspected to have caused the Magellan satellite to act erratically in its
orbit around Venus [Sellers, 1994].

Spacecraft Disorientation

Many spacecraft use star sensors to provide accurate pointing. Particles emitted by the
Sun, along with those of cosmic origin, can impact star sensors and provide false
readings. This can lead to degraded pointing or a loss of attitude control. Extreme cases
of a loss of attitude control may lead to a loss of the mission life since batteries may
discharge beyond their designed specifications and sensitive equipment may be exposed
directly to the Sun or to cold space for too long [Worden, 1996]. Other satellites that use
geomagnetically stabilised attitude pointing routines can experience pointing problems
during intense geomagnetic storming and magnetic reconnection events.

Surface Degradation

The space environment produced by the Sun can also have significant effects on surface
coatings of some spacecraft. In the Earth's upper atmosphere, the Sun causes oxygen
molecules to breakdown into oxygen atoms. Impact of these atoms on spacecraft surfaces
causes an effective oxidising reaction that is similar to rusting [Sellers, 1994, p. 68].
Another phenomenon is experienced by spacecraft which fly through the auroral regions.
The increased flux of high speed particles can cause a "sand blasting effect" on spacecraft
coatings and external sensors [Sellers, 1994, p.74]. Finally, extreme doses of ultraviolet
radiation are experienced during a satellite's lifetime which result in degradation of the
spacecraft's surface coatings and solar photovoltaic cells [Sellers, 1994, p. 71].

92 • Ra: The Sun for Science and Humanity

Magnetopause Crossings

Nominally, the Earth's magnetosphere provides a protective barrier from interplanetary
space. The Earth s magnetopause is the equilibrium barrier between the Earth's
magnetosphere and the Sun's solar wind [see section 4.4.2]. Between the Sun and the
Earth, the magnetopause usually provides shielding from the solar wind out to
approximately 10 Earth radii. However, the magnetopause can be compressed.
Occasionally, satellites at geosynchronous altitudes (6.6 Earth radii) will cross the
compressed magnetopause and be exposed directly to the solar wind. This increased flux
of particles, protons and high-energy electromagnetic radiation can create problems
within spacecraft since most are not engineered to withstand direct solar wind [Worden

4.5.3.2 Effects on Terrestrial Technology

The Sun can disrupt many terrestrial technological systems, especially the ones with
electromagnetic components. Some of the most prevalent phenomena directly linked to
the Sun that have effects on terrestrial technology are discussed below:

Geomagnetically Induced Current (GIC)

The occurrences of solar flares, and prominences on the Sun changes the magnetic field
lines in the solar wind emanating from the Sun. When this solar wind hits the Earth, it
distorts the natural geomagnetic field lines of the Earth by greatly compressing the field
lines.

As any change in the magnetic field induces current in a conductor, the changes in the
geomagnetic field lines, commonly referred to as geomagnetic storm, also induce current
in conducting materials on the Earth. This type of induced current is known as the GIC.

The GIC is most prevalent in high latitude countries like Canada and Sweden, because
significant geomagnetic storms take place mostly near the North Pole, or the South Pole;
and usually, in these places, the long power lines take the place of conductors carrying
the GIC. The effects of the GIC can range from small irregularities in voltage output to
large saturation of current in transformers, saturation to such an extent that sometimes
the transformers have been known to bum up.

An example of technology affected by the GIC is electrical power transmission line. On
March 13, 1989, in Montreal, Canada, due to the GIC some six million people were left
without electrical power for 9 hours, and quite a few elsewhere were left without power
for a few days. The financial loss to the power company was estimated to be over ten
million U.S. dollars. During this time of geomagnetic storm, some cities in the northern
part of the U.S., and Sweden were also left without power [Campbell, 1995],

Another example of technology affected is the transnational petroleum pipelines made of
conducting materials. The geomagnetically induced current in the pipelines can lead to
erroneous readings in the flow meters of the pipes, which usually results in high
corrosion rates in the pipelines.

In addition to its effects on power transmission lines, and petroleum pipelines, the GIC
also affects telecommunications cables, precision instruments, manufacturing equipment
and computers [ARINC, 1996]. n v

Fluctuating Geomagnetic Field (FGF)

Like geomagnetically induced currents, fluctuating geomagnetic field is also caused by
changes in the solar wind. An example of affected technology is the scientific equipment
used for geological explorations. Geological surveyors use magnetometers to detect
minute changes in the Earth's magnetic field to locate oil, gas, and other mineral deposits.
This type of exploration can be impossible during periods of high solar activity due to
fluctuating geomagnetic field. Another example of affected technology is magnetic
compass used for air and sea navigation. In addition to its effect on equipment used for
geological exploration and the magnetic compass, fluctuating geomagnetic field also
affects precision instruments, manufacturing equipment, and computers [ARINC, 1996].

4.5.3.3 Effects of the Sun on Radio Links and Propagation

The Sun can also have severe effects on radio propagation. Problems have been
documented with satellite and ground communications as well as radar propagation and
the GPS navigation signal.

Satellite Communications

Satellite communications experience radio frequency interference when a radio energy
burst from a solar flare occurs at the right frequency and when the receiver is in the field
of view of the Sun. The knowledge of such radio bursts enables the operator to
determine the source of interference [Worden, 1996]. The IPS Culgoora Solar
Observatory uses instruments to monitor solar radio bursts in the frequency range of 18-
1800 MHz. Radio bursts are often emitted during solar activity in addition to other
elements which cause the disturbances. Hence, their monitoring enables the prediction of
other following emissions and the disturbances that may result [Culgoora Solar
Radiospectrograph, IPS Radio & Space Services, WWW].

A similar geometry related effect called solar conjunction occurs when the Sun is aligned
with the spacecraft as seen from the Earth station. This problem does not require a solar
flare to be in progress but is much more pronounced at solar maxima when the Sun is a
strong background radio emitter. The spacecraft's orbit will determine the number and
duration of solar conjunctions. The level of interference depends upon a number of
factors including the antenna radiation pattern, the receiver bandwidth, the acceptable
signal to noise ratio and the Sun's temperature that is a function of the frequency used
and the solar activity [Solar Interference to Satellite Communications, IPS Radio & Space
Services, WWW]. For geostationary satellites, solar conjunctions will occur around the
March and September equinox due to simple geometrical considerations [Maral and
Bousquet, 1993] and calculations of antenna noise temperature increase can also be found
in this reference. Similarly, solar conjunction in the case of aircraft can cause jamming of
air-control radio frequencies.

Plasma density instabilities at the F2-region altitude of the ionosphere lead to the
ionospheric scintillation effect. Through rapid, random variation in signal amplitude,
phase and/or polarisation this will cause strong amplitude fading and phase fluctuation
to most frequencies currently used by satellites, namely UHF (0.3-3 Ghz) up to C-Band at
the high frequency end [Kivelson and Russel, 1995]. Different mechanisms will cause
scintillation at high latitude and equatorial regions and resulting in some frequencies
being more affected in a region [Secan, 1996].

94 • Ra: The Sun for Science and Humanity

Ground Communications

HF or short-wave (3 to 30 MHz) radio communications systems traditionally use the
ionosphere to "bounce off" and get extended transmission ranges. However, increased
X-rays emission during solar flares increase the D-region's electron density which in turn
can absorb HF signals. This leads to what is referred to as short-wave fade events.
Moreover, the variation of the solar ultra-violet flux during the solar cycle results in
changes in the range of frequencies available to HF communications [The Diverse Effects
of Solar Events, IPS Radio & Space Services, WWW], LF and VLF communications are
ducted by the ionosphere, thus sudden changes to the ionosphere can produce phase
anomalies in these communications and range errors on navigation systems using these
frequencies. ' °

Radar Systems

The enhanced, irregular ionospheric ionisation can produce a phenomenon called "Radar
Aurora which is an abnormal radar signal back-scatter on polar-looking radars. The
impacts include increased clutter and target masking, inaccurate target locations, and
even false target or missile launch detection [Worden, 1996], RFI also affects missile
detection or spacetrack radar.

Another effect of the ionosphere is the refraction and delay of UHF/SHF radio waves
from missile detection and spacetrack radars. This leads to target bearing and range
errors that can be compensated for based on the expected ionospheric Total Electron
Content (TEC). TEC values, however, can be invalidated by individual solar and
geophysical events.

NAVSTAR Global Positioning System (GPS)

The severe plasma density instabilities described above can also cause errors in
individual GPS navigation signals. The scintillating effect of these plasma patches
produces transmission path delays between satellites and receivers. Because the system
measures signal time delays, any phase variation will cause a time delay and will
introduce an error in the navigation solution. As of now, no conclusive studies have been
completed that characterise potential error sizes in GPS due to ionospheric scintillation
[Bainum, 1996]. Another potential problem with the GPS system is signal fade. Each
GPS receiver is designed with a TEC gradient threshold. The edges of plasma patches are
characterised with sharp TEC gradients. Sustained gradients will cause users to lose lock
on the GPS signal [Bainum, 1996]. Ionospheric scintillation of GPS is a regional
phenomenon and seems to only be observable at the poles [Bainum, 1996] and at the
Earth's magnetic equator [National Space Weather Program, 1995],

4.6 The Sun as a Resource

A way to look at the Sun is to view it as a resource. From an applications point of view
this enables one to recognise a wide variety of applications related to that Sun. Four
different types of resources are identified and described.

4.6.1 The Sun as an Energy Resource

The Sun has been the main source of energy for our planet since the beginnings of time.
Plants depend on sunlight to produce oxygen without which we could not survive.
Humans have devised ways to increase the benefits of sunlight, ranging from its use in

Our View of the Sun • 95

the production of salt from sea water to solar cells for domestic and industrial use. In
space, the Sun is the main energy provider for spacecraft.

Energy

► Education

Entertainment

^ Disposal

Earth

Space

Fig. 4.15 The Sun viewed as a resource.

Solar energy on Earth

As traditional energy resources like coal, oil and gas are becoming scarce and have major
environmental impacts, alternative sources of energy are becoming more and more
important. Solar energy is one of the most promising sources of energy. Energy can be
generated using solar cells or heat-exchangers. Focused solar energy can be used for high
temperature manufacturing uses. Significant potential energy savings could evolve from
efficient heat/light technological infrastructures implemented in buildings and
transportation media.

Solar energy in space

In space the Sun provides the main source of energy for spacecraft through the use of
solar cells that provide the electric power. A major problem with solar cells is
degeneration due to radiation. Besides that efficiencies are relatively low. New
developments in solar cells technology focus on increasing efficiency, decreasing
degeneration and methods for regeneration of solar cells. For propulsion purposes solar
sails offer a new way to utilise the Sun's energy [see section 6.4.3].

More futuristic plans involve collecting solar energy in space and sending it down for use
on Earth. The basic technology to perform such a task is available, however the market
for this kind of energy still does not exist [ISU, 1992].

4.6.2 The Sun as an Education Resource

The Sun is an education resource in the way that it has a large influence on our daily life.
Being the closest star, the Sun provides us with an excellent study object for research into
the mechanisms that make it work. See section 8.6.1 for further discussion.

4.6.3 The Sun as an Entertainment Resource

With the auroral lights, the Sun provides us with one of the most impressive features of
nature. Given the attractiveness of auroras a business opportunity might exists for their

96 • Ra: The Sun for Science and Humanity

accurate prediction. Assuming that an aurora could be predicted with an accuracy of

90 /o or better, tours could be organised to places where the aurora is visible, either on
Earth or in the sky.

Helioseismological oscillations (i.e., sunquakes), when transformed to the sound
spectrum might provide entertainment to those who want to be closer to nature.
Listening to the sounds of the Sun might very well fit in with current New Age trends.
Remember, people are already listening to whales and forests.

4.6.4 The Sun as a Disposal Resource

Due to its high temperature, the Sun is able to permanently dispose of anything by
breaking it down to protons and electrons. During solar storms, the increased solar wind
disposes of some of the space debris in low earth orbit.

The safe disposal of nuclear waste is one of the most important waste problems humanity
is facing. Nuclear waste takes thousands of years to degrade to benign matter. Nuclear
waste could be permanently disposed of by shooting it into the Sun. The obvious
problem with this solution is that the nuclear waste will have to be launched in orbit. A
launch failure of a launcher carrying nuclear waste would have severe local
environmental impacts. Because of this the political willingness to even consider the
possibility is very low.

Fig. 4.16 Space tourism, the next step (Courtesy of H. M. Rehorst).

Our Vipw nf fho Qnn • 07

Chapter 5

Objectives & Requirements

In this chapter we put forward the objectives deemed to be most important to the
trategic Framework. The first section discusses the scientific objectives, and the second
section discusses objectives related to applications. Next these objectives are compared to
the objectives of past, current, and planned solar missions and are linked to the Strategic
Framework. Finally we offer recommendations for Near-, Mid-, and Far-Term mission
requirements.

5.1 Science Objectives and Priorities in the Ra Strategic
Framework

To guide the development of the Ra Strategic Framework, it is essential that the scientific
objectives for such a programme be clearly defined. Several related lists have been
published, either in scientific literature or by agencies. Most of these refer to single
campaigns (e.g. FIRE) or a programme of missions (e.g. Solar Connections). Of course
many published scientific objectives have already been met, either fully by a completed
mission, or partially by current missions such as SOHO. We compiled our own list of
objectives based on our view of the situation in August 1996 and advice from a number of

visiting lecturers at ISU. Their input helped revise our original set of objectives and focus
them more precisely.

In particular the importance of stereoscopic imaging was stressed, as well as observations
at high spatial, spectral, and temporal resolutions, and long duration to provide
information on physical processes such as magnetic reconnection.

The objectives listed in section 5.1 apply to the whole Ra Strategic Framework, and as
such can not apply to (or be achieved by) a single mission. They are to be used in
conjunction with other objectives (such as applications and policy objectives) to guide the
development of actual missions. 6

Science priorities are always challenging subjects because scientists' opinions differ. For
the Ra project we have chosen our own priorities and we defend them by references to
scientific literature. The listing of the objectives does not imply the order or priorities of
importance.

5.1.1 Primary Objectives:

Many, if not most of the processes happening in, on and around the Sun are poorly
understood, such as the neutrino problem, the origin of the Sun's magnetic field and its
connection to differential rotation, and the solar cycle.

However, for determining how important a specific scientific objective is, we chose as a
criterion its relevance to Earth. This goes partly hand-in-hand with the application-type
and Earth-relevant objectives. To come up with better space environment predictions, we
need to understand the physics behind the phenomena that trigger magnetic storms. Seen
from a longer-term perspective we are even more worried about the Sun's influence on
potential climate changes. Thus we divided our primary objectives into exactly these two

categories.

To understand the physical processes leading the Sun t o emit plasma structures and high
energy particles that are potential threats to humans and technology.

This automatically leads to the following issues to be addressed.

• What is the heating mechanism of the corona?

• What leads to the formation of coronal holes?

• From where does the slow solar wind emerge?

• How intimately is the fast solar wind related to coronal holes?

• What are the causes for and underlying physical principles of solar flares?

• What are the causes of the acceleration of particles to very high energies?

• What leads the corona to release coronal mass ejections?

• How do the different types of coronal mass ejections propagate in the
interplanetary medium?

To answer these questions it is essential both to develop new observational techniques,
such as stereoscopic imaging of the corona, and to improve theoretical models.

To understand the physical processes which may lead the Sun to influence our climate.

This automatically leads to the following questions:

• What causes the solar "constant" to change?

• What are the long-term variations in the solar constant?

• To what extent do variations in the solar constant influence the Earth's
climate?

5.1.2 Secondary Objectives:

We determined the following objectives (not directly related to the Sun s influence on
Earth) to be secondary:

100 • Ra: The Sun for Science and Humanity

• To determine the cause of the solar cycle.

• To determine what causes the solar constant to change.

• To investigate the origin of the Sun's magnetic field and its connection to
differential rotation.

• To determine the internal state of the Sun by measuring the higher

harmonics of its gravitational field. " °

• To determine the internal state of the Sun by means of helioseismology.

• To test general relativity by using the Sun's gravitational field.

• To measure the abundance of galactic cosmic rays in the Sun's vicinity.

• To solve the neutrino problem.

The first three secondary objectives are very closely connected to the primary objectives-
however, we chose to make the distinction as above. On the one hand we placed
emphasis on the effects that a changing solar constant might have on Earth, as opposed to
its cause, which is a phenomenon related to the interior of the Sun. Similarly, we did not
ask for the origin of the magnetic field, instead placing emphasis on its effects.

5.2 Applications Objectives and Priorities in the Ra Strategic
Framework

To keep the mission objectives input to the Ra Strategic Framework as comprehensive as
possible, a broad view of the possible nature of missions to the Sun was taken. This view
went beyond the traditional science-only missions view and included the possibility of
applications-focused missions. From an applications perspective the following three
goals were adopted to derive inputs for the Strategic Framework:

• identify and investigate solar-terrestrial missions dedicated to a particular
application,

• identify and investigate application spin-offs from science missions, and

• identify and investigate future applications that require technology

development, 7

all for the benefit of humanity and commerce.

5.2.1 Applications Needs and Opportunities

To assess the needs and opportunities for solar-terrestrial related applications it is helpful
to consider the Sun as either a threat [see detailed description of section 4.5] or as a
resource [see overview of section 4.6]. Since utilising the Sun as a resource was the focus
of a previous ISU report [ISU, 1992] it was decided to focus on responding to the Sun as a

threat. Two different categories of a response to a threat are possible:

• either, eliminate the threat by preventing it from occurring, by deflecting it,
or by continuously protecting your system from the threat,

• or, mitigate the threat by predicting its impact and taking appropriate
safeguard actions.

Based on our current state of knowledge concerning the threats outlined in section 4.5,
threat elimination was not considered feasible although opportunities for protection
technology development are numerous (e.g. thermal shielding, radiation hardening,
discharging techniques, etc.). These technology oriented issues are explored in chapter 6.

Objectives & Requirements • 101

5.2.2 Applications Focus

The chosen applications focus was therefore on mitigating the harmful effects of the Sun
by predicting their occurrence and making it possible to temporarily safeguard systems,
i.e. Solar Threat Monitoring and Early Warning. In the Near-Term this would include
increasing the awareness of solar event impacts and improving the use of current
resources [sections 8.6 and 8.4], in the Mid-Term this would possibly include applications
oriented science mission enhancements and/or the implementation of a dedicated early
warning system [section 9.2], and in the Far-Term this would include future applications
requiring technology development [section 10.1.2] plus a permanent, world wide
prediction and warning system.

To justify this focus we made a survey of the existing solar threat monitoring and early
warning systems and we found that no dedicated system currently exists [see Appendix
E: Existing and Proposed Early Warning Systems]. The current state of the art is
opportunistic in terms of its acquired measurements and the result is probabilistic, not
unlike Earth weather forecasting in the past! This need not be the case given advances in
our understanding of the triggering mechanisms of magnetic storms and advances in
sensor technology. The goal of section 9.2 which explores different options for a
dedicated early warning system is to define a system that will make solar threat
monitoring and early warning more deterministic and far less probabilistic.

5.3 Mission-Objectives Analysis

The aim of this paragraph is to analyse the current scientific and application objectives
discussed in sections 5.1 and 5.2 and perform a comparative analysis among the
objectives that have been defined for the past, current, and planned international solar
missions. Space research can provide us with more comprehensive information needed
for understanding, predicting and monitoring solar activities for the benefits of
humankind. The measurements performed by each mission to fulfil its scientific and
application objectives are categorised as depicted in Figure 5.1.

C/D

h-

Z

LU

2

LU

DC

=>

C/D

<

LU

i—► Electric

•FIELDS —[*■ Magnetic

Gravitational

Acoustic

- WAVES —M* Electromagnetic
!-► Gravitational

“ ■ T’ * " "

Intensity

Gradient

Power

Polarisation

•PLASMA H

Particles

,• : : /U

!-► Electrons

' •iii*;') s’ ; •:* ' ’

-► Protons

-►Ions

' •■'-,y : .v '■ '

-►Others

mllilW

^pligilySs':

X....

: v: v

m t? rj ■ m i:

Fluid

Temperature
v..- _ Heat Flux

, ■ s; :: - Density

■IMAGING

WWM

Whole Electromagnetic Spectrum

Fig. 5.1 Categorisation of measurements.

102 • Ra: The Sun for Science and Humanity

5.3.1 International Missions Objectives Background

This paragraph will describe and analyse the specific objectives of the past, current and
planned solar missions (see Tables 5.1, 5.3. and 5.5). In the measurements tables [see
Tables 5.2, 5.4 and 5.6], the regions in space where the spacecraft have been collecting
data are divided in three [see Figure 5.2]:

Region 1: Close to the Earth, up to 30 R E ;

Region 2: Intermediate region, from 30 R E to 30 R s to the Sun;

Region 3: Near the Sun, closer than 30 R s from the Sun.

5.3.1.1 Past Missions

The period 1962-1980 has been arbitrarily chosen, even if some spacecraft launched at
that time are still in operation today. The missions during this 18 year period have
covered various objectives, have been launched on a variety of trajectories and have been
implemented through a number of significantly different collaborative agreements The
scientific objectives of these spacecraft seem to have been global, no mission was specially
designed to one specific objective. On the contrary, every mission carried experiments
and instruments covering multiple scientific objectives. In the survey and assessment of
past missions, there is no evidence of any substantial or direct interest in applications
based either on the availability of solar related environmental information or Sun-Earth
interaction. The main emphasis has clearly been on improvement of our knowledge of
the Sun and interplanetary medium and solar system/Sun related environmental
information, to prepare manned space missions and to cope with disturbances to Earth-
orbiting artificial satellites. The national programs (US and USSR) are more numerous
than the international ones. However, there were some bilateral partnerships between
countries (USA / Germany, USSR / France) or between agencies (NASA / ESA). The
trajectories of the spacecraft were very different. Some were in low Earth orbits, others
were in intermediate Earth orbits. It is in this period that the mission to date closest to
the Sun (Helios) was successfully conducted. Several interplanetary spacecraft were
carrying instruments to study the Sun even from high latitude (Ulysses).

OblRftlVP8 Rpnniromontt- * in'!?

5.3.1.2 Current Missions

There is a 10-year gap between current missions and past missions. Solar Max was
launched in 1980 and Ulysses was launched in 1990. In the intermediate period only a
few Prognoz spacecraft were launched. Why this gap? We assume the scientific
community has been analysing the data gathered by the previous missions while at the
same time preparing combined, continuous and co-ordinated Sun s study programs,
within the ISTP or IACG organisations. Objectives covered the whole range of scientific
fields of interest at this time. More missions focused on particular fields, some of the
most important being the corona, solar flares and the CMEs. The interest for Sun/Earth
interaction increased during this period and some missions are more focused on these
objectives. The majority of the trajectories and final orbits were near Earth, at low or
intermediate altitudes, with only Ulysses orbiting over the solar poles and no spacecraft
at an approach distance closer than 64 R s .

5.3.1.3 Planned Missions

The planned missions appear in two different types: the ones that are already scheduled
with a definite launch window and very precise characteristics, and the ones that are still
in the approval cycle. Among the last ones we find the missions designed to complete
measurements of previous missions, in particular those co-ordinated through the ISTP.
Sun/Earth interaction studies have an important role in the forthcoming period and
environmental effects of solar activity are more precisely assessed. The corona is the
centre of interest in almost all planned missions and for the first time plans have been
established to send spacecraft closer to the Sun to make measurements from very small
distances in high temperature environments. Important programs launched in the
beginning of the 90's are about to reach their completion, and in the present schedule
there are no foreseen replacements. At the same time the Cluster constellation was lost in
a launch vehicle failure representing a significant set back in the program. Are we going
to have another empty decade such like in 1980? From the co-operation point of view, we
do not find the same strategy adopted as in the previous period; no ambitious joint
program such as SOHO, CLUSTER or ULYSSES exists; only some bilateral or trilateral
project is being considered. However, CLUSTER recovery options are being studied and
evaluated by ESA and the science community.

Table 5.1 Past Missions: General Objectives.

1962

1971

1972

1972 |

1973 |

1974

1977

1977

1978

Missions

OSO

(8 s/c)

SOLRAD

(3 s/c)

Pioneer

(11 s/c)

m

»

HELIOS

(2 s/c)

Voyager

(2 s/c)

SIGNE 3

(1 s/c)

boiar

Max
(1 s/c)

Primary Sci objs

■

■ ■

HHliH

L — ■■

- Solar Corona

■■■

•t ‘ *•*%; -v

- Solar wind

- Farth/Sun

Hi

mmm

BBS

■

1 KT J*f*l il* L 1

- Innpr Sun’s Physics

- Gravitation

|P ■

- Cosmic ravs

I1ETS253

fi

!■■■

- Threat apps

- Resource apps

ina • Ra- The Sun for Science and Humanity

Table 5.2 Past Missions: Measurements.

Table 5.3 Current Missions: General Objectives.

Objectives & Rermi rpmpnk • 1 HR

Table 5.4 Current Missions: Measurements.

Table 5.5 Planned Missions: General Objectives.

106 • Ra: The Sun for Science and Humanity

Table 5.6 Planned Missions: Measurements.

5.3.2 Comparative Analysis

The first conclusion of the analysis is that no long-term strategy has been adopted to
define the solar missions that have been flown or developed so far. International co¬
operation has been promoted only recently so that many similar missions have been
conceived by different countries without there being any correlation. The number of
necessary missions can eventually be reduced and the on-board instrumentation can be
optimised if a comparative analysis is performed on the measurements.

Four other main observations can be made by analysing the past, current and planned
missions:

1. The corona has been studied from 1962 up to now by 11 out of 20 past and
current missions; while 7 out of the 11 planned missions plan to collect
more data. Despite this fact the corona remains to be one of the most
mysterious regions of the Sun. From a scientific point of view we conclude
that we need measurements different from those made up to now, from
different observation locations (L4), from closer orbits to the Sun (maybe
suicide probes) and by different means (3D imaging, stereo imaging).

2. ISTP programs are today giving us very good data on the influence of the
Sun on terrestrial environment. However GEOTAIL will end its mission in
1996, Wind and SOHO in 1997 and Polar and INTERBALL in 1998. Even if
their lifetime will be extended, no additional missions are scheduled to
replace them during the next decade using a similar international co¬
operation. Cluster was an important part of the ISTP and its launch has

Objectives & Requirements • 107

failed so valuable data are missing today to achieve the goals of co¬
ordinated observation for the ISTP.

3. Up to now only a few spacecraft have been dedicated to study inner Sun
physics and none are planned up to 2004. We assume it is because a lot of
data on this subject can be gathered from Earth or from non-dedicated
spacecraft making remote measurements of gravitational or acoustic waves.
However, even if inner Sun physics is a secondary objective for scientists
maybe it should be emphasised more in the Mid- or Far-Term programs.

4. Applications are quite absent of all past, current and planned missions,
even though indirectly data are being gathered by existing spacecraft
(WIND, SOHO) and are used for monitoring the space environment and
forecasting Sun / Earth interaction. Today the need for such forecasts is
increasing. Private space companies, governmental agencies and even
human every day life are more and more concerned about it. Such an
objective would likely get a large approval consensus among decisional
entities.

5.4 Scenarios

This section provides a technical link between the analysis presented in the previous
section and the Strategic Framework. It depicts the multiple dimensions of a Sun
exploration mission, and lists the options available today or in the Near-, Mid- or Far-
Term, if any change is foreseen. This allows to match the means to the needs.

5.4.1 Needs and Measurements.

A conservative, step-by-step approach, without new missions is necessary in the Near-
Term. Mid-Term is concerned with low-risk applications offering a material benefit to
the community. Far-Term addresses more ambitious questions about the corona and
inner solar physics, taking advantage of new technologies. Viewed today as 'enabling',
these technologies should become mature in the 15-25 year Far-Term time frame.

5.4.2 Spacecraft Fleet and Trajectory

Increasing the number of spacecraft in a mission allows stereoscopic and/or time-spread
measurements, helping the analysis of Sun processes. Miniaturisation could help to
conserve total mission mass, avoiding launcher penalty. This will depend on the
improvement in mass and volume of instruments, electronics and thermal shielding, and
likely is a Far-Term opportunity. In the Mid-Term, 'a few' spacecraft per mission seem
preferable, helping to master intercommunication and control questions for later
constellation missions. Size is affected by propellant mass, i.e. trajectory, mission
duration and propulsion technology. Chemical propulsion gives too low speed levels.
This imposes to use gravity assists, a long process that suffers from the low solar energy
available for on-board power (Jupiter) and suffers from long link distances.

Getting 'closer to the Sun', and 'more often', are two scientific repeated requests, that are
expensive and long to achieve with chemical propulsion. However two alternatives look
promising: first electric propulsion and then solar sails. Electric propulsion is currently
planned for demonstration in the US New Millennium program and offers much greater
jet velocities allowing closer access to the Sun. Because of its relative novelty, it is a Mid-
Term to Far-Term option. Solar sails offer similar advantages to electric propulsion but
are considered as more unconventional. Deployment and survivability close to the Sun
appear as challenges, although the capability of changing orbit inclination is attractive for

• Ra‘ Thp Sun for Science and Humanity

high latitude measurements and mapping. This makes solar sails attractive for Far-Term
constellations.

Table 5.7 Needs and Measurements.

Field

Options

Trade-Off

Solar Science

Corona:

cause of heating, cause of CME,
dust at <0.3 AU, holes,
cause of flares, EM field.

Solar Wind: origin and process, polar wind.
Sun-Earth Interaction: Earth weather, effect of
Sun on Earth magnetosphere, magneto-iono-
atmospherics.

Secondary items: sunspots and their EM field,
solar 'constant', Sun gravity field, seismology,
cosmic rays near Sun.

In situ vs. remote sensing.
Ecliptic vs. inclined
trajectories.

Field or particle
instruments.

Applications

GIC prevention, power line and sat protection,
EVA protection, public and leisure, power
generation, energy-efficient technologies.

Instruments

Philosophy:

in-situ,

remote sensing: EM spectrum through solar
layers.

single measure vs. imager.

Particle: plasma analyser, energetic particles
detector, dust detector.

Field: magnetometer, gravity gradiometer.
Waves: visible, IR, microwave, X, K-band.

White light coronograph, EUV telescope

In-situ is more dangerous
to spacecraft.

Power, atmospheric
attenuation

Data quantity and transfer
rate.

EUV and microwave allow
to relate corona with
photosphere.

Recommended

Requirements

Near Term:

- continue existing missions,

- use other observation means (observatories,
mil sats),

- improve data management and distribution.

Medium Term:

- develop applications related to Earth
protection.

- develop scientific missions on Sun/Earth
interaction.

- improve international co-operation.

- set up long duration observation
programmes.

- optimise Instruments suites per s/c.

- develop constellations for multiple
measurements.

Far Term

- address solar physics.

- develop in-situ missions and 3D
measurements

- explore space collection of solar energy.

Table 5.8 Spacecraft Fleet and Trajectory

Field

Options

Trade-Off

Spacecraft

Number

Single,
a few,

constellation.

suicide probe

3D and multiple
measurements, series
production effects,
risk spreading, launcher
size.

Shorter total duration.

Deeper exploration.

Trajectories

Orbit:

around Earth: LEO, synchronous,
around Sun: circular: at 1 AU, 30 R^..

elliptic: at 30 R* or more, 4 R*...

Remote sensing.

In-situ sensing. High
velocity.

Duration, comm., heat.
Shortly close to Sun.

Direct

Gravity Assist at Jupiter: out of ecliptic or for
ecliptic circularisation.

Resonant Venus GA perihelion at 0.25 AU,
inclined at 20°.

Helicoidal: see ion thrust or solar sail.

Too costly, especially out

of ecliptic

long.

Propulsion

Solid Chemical has the lowest jet speed and
can not be switched.

Liquid Chemical is limited to 5 km/s jet
velocity.

Electric offers very high exit velocity but very
low thrust.

Solar Sails.

Can not propel fast enough
for solar orbit.

Needs demonstration.

Flying in New Millennium.

Needs robustness to
survive.

5.4.3 Environment and Subsystems

The environmental constraints mainly concern the extremely wide variations of
parameters to be coped with by the spacecraft. Jupiter assists imply low solar energy for
on-board power, low temperature and long flight time and communication distances.
Proximity of the Sun involves thermal shielding and signal/noise separation issues.
Earth-Sun celestial mechanics imposes very high spacecraft speeds, exceeding current
capabilities.

Some subsystems technologies should alleviate these issues. Carbon/carbon is the
shielding material of choice, up to about 4 R$. Cost issues might however restrict Mid-
Term mission to trajectories further from the Sun. In the Far-Term however, high
temperature electronics and optical communications should make more affordable the
closer solar orbits desired for in situ observation.

110 • Ra: The Sun for Science and Humanity

Table 5.9 Environment and Subsystems

Environ¬

ment

Transmissions: Sun-Earth line noise.

Heat: current heat shield up to 4 R s .

Outgassing: from s/c, might corrupt
measurements.

Particles: solar flare first result in protons that
are dangerous for electronics, and then in heavy
ions causing electronics upsets. High speed
particles might be catastrophic.

Radiation is significant in planetary
magnetospheres.

Electrostatic charging is induced by solar
plasma. Discharging might damage
subsystems.

Magnetic induction might cause perturbation
torques and blur measurements.

Heat

Protection

Carbon/Carbon,

Ceramics,

Convenient,

emissivity/absorptivity.
Brittle, UV sensitive.

Refractory Alloys

Mass loss.

High Temp Composites

Relatively low tempera¬
ture.

Communi¬

cations

Outer Corona will affect transmission
amplitude and phase.

Data Storage can relieve transmission issues
close to the Sun.

Distance affects communication sizing (Jupiter).
Sun-S/C Separation is negligible below 4 R s .

Data rates lead to consider SHF/EHF and X, Ka-
bands.

Microwave transmission relies on frequency
windows in the ionosphere.

Depends on storage
duration

Need to develop high
frequency transponders,
Ka-band stations.

Optical links offer greater data rates due to
greater frequency. They avoid scintillation from
corona and solar wind.

Electronics Temperature:

current electronics operates up to 65°C.

SOI, silicon on insulator, operates up to 300°C,

SiC electronics operates up to 600°C.

Allows coherent light
detection, discarding Sun
noise, but is attenuated by
atmosphere.

Needs new receiving
telescopes, better in orbit
(Earth or libration point).

Power

Large variations
temperatures and in solar
flux (3% of Earth level at
Jupiter).

Solar Arrays: classical or with concentrator.

Concentrator is 1 /2 present
cost, better hardened and
uses higher voltage.

Fuel cells, electrolysing water

Nuclear Generator

Any power, but heavy and
delicate.

More compact and lighter
than solar arrays, but
difficult to launch and less
efficient. Policy restriction.

RTG, radioisotope thermoelectric generator.

Electrodynamic Tethers

Solar Heat Converters

Expensive, creates high
radiation and heat.

Wire needs deployment and
insulation.

Bimetallics are 5-7%
efficient, thermionics are 20
% efficient.

Objectives & Remit remonfc m

5.5 Recommendations on Requirements

Based on the analysis in sections 5.3 and 5.4 these are the recommendations for Near-,
Mid-, and Far-Term mission requirements.

5.5.1 Near-Term Missions Recommendations

In the Near-Term, in order to get, in the most cost-effective way, the data necessary to
fulfil the current scientific and application objectives defined in sections 5.1 and 5.2 we
recommend:

to focus on the solar missions under development at the moment that do not
require any particular technology development and co-ordinate them.

• to look for any other potential sources of data about the Sun/Earth
interaction to be used for the benefit of the Earth environment (military
satellites, and observatories world wide).

• to improve international data availability and management.

5.5.2 Mid-Term Missions Recommendations

For the Mid-Term missions the solar science benefits should be the main goal to be
achieved. Therefore we recommend:

• to focus mainly on the fulfilment of the application objectives related to the
Sun as a threat (solar weather monitoring and early warning), as this would
minimise economic damage to industrial equipment,

• to focus on the primary scientific objectives related to the effects of the Sun
on the Earth,

• to promote world wide international organisation co-operation, paying
particularly attention to developing countries,

• to assure continuity of observations on a long-term basis,

• to focus on missions related to region 1 and 2 (Distance from the Sun greater
than 30 R s to the Sun).

No particular time correlation in measurements being required for these missions, each
spacecraft should be oriented to a specific measurement category (fields, waves, plasma,
images) and the number of objectives to be fulfilled should be optimised on a
measurements based criteria. Following this approach the spacecraft structure can be
optimised in relation to the type of measurements to be performed, resulting in a reduced
weight reduced interference among instruments, increased overall performance, and
lower costs. Small spacecraft constellations, using possibly a common bus are suggested.
Daily monitoring would generate information useful for scientific analysis and solar
model improvements. No technological leap would be required, but several
improvements could ^spin-ofF for later missions: pilot use of electric propulsion,
spacecraft to spacecraft communication, smaller scale electronics and self-healing
software.

5.5.3 Far-Term Missions Recommendations

For the Far-Term missions requirements we recommend:

• to focus on the fulfilment of the application objectives related to the Sun as
both a source and a threat,

• to focus on the fulfilment of the scientific objectives related to solar physics

and theory. J

The fulfilment of the scientific objectives requires specific in situ measurements. Time
correlation measurements being the key for most of those observations, a mission design
should be based on the use of multiple spacecraft in the same spatial region taking
simultaneous measurements. Each spacecraft should be optimised for a particular
measurement category taking advantage of the related optimisation design experience
gained m the Mid-Term. Technological improvements should make deep exploration
and m-situ multiple-latitude mapping missions, able to gather data on macro and micro
solar processes, affordable thus allowing revision of current solar physics understanding.
This extensive collection of information should help to discover solar physical principles
that remain unknown today. This should help to advance the sciences of matter and their
applications such as electronics and computing.

The enabling technologies would be a combination of electric propulsion and or solar
sails, robust solar arrays or solar heat converters, high temperature electronics, optical
communication with Earth-orbiting relay spacecraft. The development of constellations
should benefit from spacecraft 'series' production, modularity of sensors, and from image
fusion with improved database management. The smaller more numerous spacecraft
would be better suited for incremental improvement and make the system more failure-
tolerant.

Chapter 6

Technology Challenges and
Issues

A mission to the Sun presents many technological challenges due to the harsh and
extreme environments that a spacecraft will encounter. The purpose of this chapter is to
document the anticipated technological challenges to the Ra missions and to provide a
menu of available technologies, including their advantages and disadvantages.

6.1 Solar Environment

The space environment is a key challenge in the design of spacecraft. For solar missions
all the different conditions experienced from the geocentric parking orbit, eventually
gravity assist near a planet, and heliocentric orbit must be addressed. This section gives a
brief introduction to the specific issues under concern for interplanetary missions,
specifically with focus on close solar approach. The interplanetary environment is in
many cases different from the Earth's atmosphere, as described in section 4.5.3.1.

6.1.1 Electromagnetic Disturbance

Communication between ground station and the spacecraft can be problematic as the Sun
emits electromagnetic noise in all radio frequency bands. The most severe case is when
the spacecraft is close to the Earth-Sun line as periodically will be the case for the
heliocentric orbits. High gain antennas are required and very narrow beam receivers
need to be used on ground.

6.1.2 Solar Infrared and Visible Radiation

The solar radiation becomes increasingly more severe when going close to the Sun. The
thermal energy must be dissipated to provide a proper operating temperature range for

the payload. Heat shields and thermal control can be designed to go as close as four solar
radii (see the Solar Probe mission [Randolph, 1995]). In our case, the heliocentric
missions with orbits down to 30 solar radii are different in the sense that the spacecraft
have less radiation, but must be designed to live for several years. Outgassing of material
from the shield must be minimised to avoid contamination of the scientific instruments.

The trajectory selection is central in the design of solar arrays, as the available power
depends on the distances to the Sun. Far away from the Sun the flux is approaching zero.
This fact is part of the reason for avoiding Jupiter gravity assist in the design of the Ra
missions. Close to Sun, the solar arrays are heated causing degraded performance.

6.1.3 Particle Radiation

High energy particle radiation can have hazardous effects on electronics. Microstructural
damage leads to degradation and possible failure. Proton radiation with energies above
30 MeV, which increases in density with solar flares, can be extremely dangerous. Single
Event Upsets are caused by heavy ions from the galactic cosmic radiation and increase in
solar wind energetic particles following solar flares. Particle radiation is a problem
anywhere in space, but more energetic particles are trapped in the magnetic fields of the
planets. The problem is therefore particularly important for periods when the spacecraft
is close to Earth, and even more serious if the spacecraft goes by Jupiter, which has
extreme radiation belts [Petrukovich et al., 1995], [Tascione, 1994].

6.1.4 Surface Charging

The electrostatic surface charging of a spacecraft when it penetrates the solar wind
plasma must be considered. A voltage potential of the spacecraft, due to photoelectric
effects, disturbs measurements of charged particles. Furthermore, discharging can cause
spurious electronic switching, breakdown of thermal coatings, and degradation of solar
cells, amplifiers, and optical sensors [Tascione, 1994]. The main contributions to charging
come from the plasma electron current, photoemission current, and thermal emission.
The current balance is very different for the environment of the Earth, other planets, and
heliospace, and must be considered individually. The most severe is the Jovian radiation
belt, where a spacecraft can charge up to tens of kV [Petrukovich et al. r 1995].

6.1.5 Deep Dielectric Charging

Deep dielectric charging is different from surface charging because it originates from 2-10
MeV electrons that penetrate deeper into the surface. This can create voltage potentials in
the internal circuitry and cause malfunction of computers, electronics, and instruments
[Tascione, 1994].

6.1.6 Dust Particles

Solid particles in the solar system originate from decaying comets, asteroid debris, and
interstellar grains penetrating the solar system [Morfill et al., 1986]. The impacts on
spacecraft are not very well known, but relative speeds above 100 km/s could be
catastrophic [Tsurutani et al, 1995]. When trajectories and orbits are determined, the
possible presence of dense dust regions should be taken into account. Dust rings may
exist around the Sun with densities 5-10 times larger than the overall dust density
[Mann,1995]. Details on the interplanetary dust cloud can be found in [Giese et al, 1986].

116 • Ra: The Sun for Science and Humanity

6.1.7 Magnetic Induction

When a spacecraft flies through a magnetic field, eddy currents can be generated in
structural parts that are not properly electrically bonded or insulated. This causes a

magnetic residual that can disturb magnetic measurements and generate disturbance
torques affecting the attitude.

6.1.8 Summary

The most significant environmental effects with impacts on interplanetary spacecraft
ave been briefly introduced. Detailed descriptions are covered in the specific sections
where the technological solutions are considered.

6.2 Payload Instrumentation

In this section we give a short description of the instrumentation developed for various
missions dealing with studies of the Sun as well as main problems and challenges which
may be encountered during the development of new instruments to meet our objectives.

6.2.1 Classification of Instruments

Two basic types of space instrumentation exist for use in interplanetary spacecraft.

• Remote sensing instruments measure the properties of photons or particles
arriving at the spacecraft from a distant point of origin.

• In situ instruments measure the properties of fields around the spacecraft
and associated waves and particles coupled to the environment
surrounding the spacecraft.

The boundary between these two definitions is somewhat blurred. For instance, in the
electromagnetic spectrum there is no clear boundary between radio waves arriving from
a distant source and electromagnetic waves coupled to the surrounding plasma. Wave-
particle duality blurs the boundary even more.

Instruments may be further classified into active and passive measurement methods
though for interplanetary missions most measurements are passive (exceptions include’
radar imaging of planetary surfaces).

For space-based solar physics the main tools of investigation are plasma instruments and
remote sensing of various layers of the solar atmosphere with the electromagnetic
spectrum. A basic plasma package consists of an electrostatic analyser for detection of
electrons, protons and ions, together with a magnetometer to establish the strength and
direction of the magnetic field to which the plasma flow is coupled. Extra information is
gained by also including sensors for electric field. Useful observations of the Sun may be
made in virtually every part of the electromagnetic spectrum. Some wavelengths mav be
observed rom the ground, but for the UV, X-ray and gamma ray parts of the spectrum it
is essential to go beyond Earth's atmosphere. Techniques associated with remote sensing
in the electromagnetic spectrum include the use of the Doppler and Zeeman effects as
well as polarisation. Spatial, spectral and temporal resolution are key parameters
together with field of view and aperture.

Tables 6.1 and 6.2 summarise the basic measurable phenomena and their associated
requirements for detection. The in situ phenomena in table 6.1 include basic plasma

properties as well as particles such as neutrons and cosmic rays. Gravitational fields can
only be sensed by tracking the spacecraft's motion.

Table 6.1 In Situ Measurement Types.

Subject of Measurement

Instrumentation

required

Science obtained

Fields &
Waves

Magnetic (B)

Fluxgate magnetometer
on boom

Basic plasma properties

Electric (E)

E-field probes on booms

Basic plasma properties

Gravitational

(g)

Low perihelion, accurate
clocks, drag-free motion,
accurate tracking

Heliodesy, General Relativity

Dust

Dust analyser
(various designs)

Interplanetary dust environment
& composition, interaction with
Sun

Particles

Electrons (e-)

Plasma analyser

Basic plasma properties

Ions

(p+. He 2 *,...)

Plasma analyser

Basic plasma properties

Neutrons (n)

Scintillation

Detection of solar neutrons
before decay (Tl/2=11 min)

Cosmic Rays
(CRs)

Galactic CRs

Energetic particle
telescope

Variation with 11 year solar
cycle

Energetic
Solar Particles
(ESPs)

Energetic particle
telescope

Origin & acceleration of ESPs

Table 6.2 shows the basic categories of available remote sensing measurements. In
addition to the electromagnetic spectrum there are other means of remotely sensing the
Sun, including the new technology of neutral particle imaging, as demonstrated for
Earth's magnetosphere on the Astrid satellite and due to fly on IMAGE [The IMAGE
Mission, NASA GSFC WWW]. Neutrinos are only practicably measured with many
tonnes of detection material down in mines on the Earth. This due to their small
interaction cross-section and the shielding necessary to exclude high energy cosmic rays.
During the Ra project we found no information to suggest that measurements of other
remotely-detectable phenomena (examples include gravitational waves or subatomic
particles other than those already mentioned) were of use in investigations of the Sun.

Table 6.2 Remote Sensing Measurements.

Instrumentation

required

Science obtained

■BH

Plasma analyser with filter & ioniser
at aperture

Charge exchange processes,
context for in situ
observations, early warning

Neutrinos

V. large scintillation chamber or
solid state detector with CR
shielding or discrimination

Fusion processes in solar
core

Radio

Radio wave propagation
(attenuation, refractive index,
Faraday rotation)

Plasma density, magnetic
field

Microwave

No immediately obvious
observations

Infrared

j IR imaging and spectrometry

[IAU, 1994]

Imaging solar disk and
interplanetary dust
distribution

Visible

White light coronograph,
desirably stereoscopic

Coronal structure, context
for in situ observations,
early warning

Ultraviolet

Spectroscopic imaging
of solar atmosphere

Temperature variations
with depth, location & time

Coronal structure, context
for in situ observations,
early warning

Gamma Ray

Collimated scintillator &
photomultiplier tube

High energy processes, e.g.
solar flares, e-e+
recombination

Phenomena not originating from the Sun itself but worthy of investigation include
galactic cosmic rays and the dust environment near the Sun. Galactic cosmic rays (in fact
high energy charged particles) are modulated in correlation with the Sun's 11-year cycle.
High solar activity reduces the influx of cosmic rays into the inner solar system.
Interplanetary dust has only been studied at distances greater than 0.3 AU from the Sun.
The dust community" has identified the dust environment in this unexplored region
close to the Sun as worthy of investigation [Mann, 1995],

6.2.2 In Situ Instruments

6.2.2.1 Introduction

The past 25 years of studies demonstrated that a continuous flux of charged particles
streams from the Sun past the planets and into interstellar space. An understanding of
the dynamics and solar sources of a continuous plasma outflow has been much more
recently acquired. Spacecraft whose trajectories take them beyond the Earth's
magnetospheric cavity are able to directly sample the charged particles flowing out from
the Sun. Such in situ measurements account for most of our understanding of the solar
wind near the plane of the Earth's orbit.

6.2.2.2 In Situ Instruments from the Solar Probe

With modern spacecraft technology, the last frontier for in situ exploration of our solar
system is the solar corona. Among the current solar probe missions, the Solar Probe
mission of USA has the most advanced in situ instruments. So we would like to adopt its
instruments to the Ra missions as a result of having made comparisons with those of

previous solar missions [see appendix C.l].

Technology Challenges & Issues • 119

Solar Wind Plasma Particle Analysers

The basic requirements for the solar wind plasma particle analyser are that the ion
instrument must be able to distinguish alpha particles from protons under all conditions
and measure complete three-dimensional velocity distributions. The basic moments of
the distributions, density, velocity and temperature, should be obtained fast enough and
accurately enough to enable Alfv£n fluctuations and MHD turbulence to be analysed.

3 - D Ion Velocity Spectrometer

This proposed design scheme is based on sensors (table 6.3) currently being built or
completed for flight programmes. The Proton Alpha Sensor is designed to define both the
geometric factor and the angular response. The Thomson Parabola Ion Analyser can define
the sensitivity and angular response, a magnetic deflection system, and an electric
deflection system with the electric field parallel to the magnetic field.

Table 6.3 Characteristics of the Sensors.

Instrument

Mass (g)

Power (W)

Proton Alpha Sensor

250

0.6 j

Thomson Parabola Analyser

300

2.0

Electronics box & connectors

525

Tilt table & electronics

1000

2.7

Ion Analyser

This instrument intended for specific studies of the ion population should measure
energy and mass/charge with time resolution of 10 s. Determination of charges is the
major objective of this instrument in order to unambiguously resolve key ion species like
oxygen and iron and their charge distributions, which through their freezing-in
temperatures may serve as plasma thermometers for the solar wind particles' source
regions in the inaccessible lower corona. Of the existing designs, such as the ones used
on Ulysses and SOHO, the latter one is to serve best on the non-spinning solar probe,
because it employs quadrupole lenses for FOV enlargements to a cone of 50 degrees
opening.

3 - D Element Velocity Spectrometer

The electron spectrometer will provide the electron velocity distribution within energy
ranges from 1 eV to 4 keV and from 2 keV to 20 keV and density range from 10 to 10 6 cm' 3 .
The detector accuracy must be high enough to ensure the precise determination of the
density, velocity vector, pressure tensor and heat flux vector of the electrons. The
proposed energy resolution is 15 % and angular resolution is 22.5°.

Magnetometer

The magnetometer must be able to measure coronal magnetic fields over a broad
dynamic range to study solar corona heating mechanisms, especially the mechanism for
solar wind acceleration. Hence a combination of sensors should be used. First there must
be fluxgate sensors with the possibility to switch between regimes of relatively weak
magnetic fields. In case of stronger magnetic fields the saturating fluxgate sensors should
be used instead. A combination of the DC magnetometer with a current probe would

allow a more complete in situ determination of local magnetic fields and current
properties. The range /resolutions of the instrument are: 1 mT/32 nT for
magnetoresistive channel and 64000 nT/2 nT, 3200 nT/0.1 nT, 256 nt/8 pT for fluxgate

However, since the magnetometer of Russian Fire Mission has better mass and power
characteristics, we would like to suggest using the Russian’s to save its consumption of
energy and reduce cost. r

Plasma Wave Experiment Package

The role of the Plasma Wave Experiment Package is to identify the various wave modes
hat comprise the turbulence spectrum existing within the extended coronal envelope and
to measure their intensities within the frequency range from 0 to 10's of MHz.

Supra thermal Particle Sensor

This insfrument is designed to study the low energy end of the solar energetic particle
population, particles accelerated at shock waves in the corona, and pick-up ions from
particles outgassing or being sputtered from interplanetary dust. It has two enerev
regimes: 20 keV and 1000 keV/charge

Solar Energetic Particle Analyser

The set of sensors included in this instrument has to be able to measure protons from 50
keV up to 50 MeV and electrons from 4 keV to 10 keV. The system must be flexible to
work in different regimes, since any SEP information while approaching the Sun to
distances closer than 0.3 AU is essentially new. It is necessary to use different sensors to

cover the broad range of energies for protons and electrons.

Detectors for Interplanetary Dust Particles

The aims of the dust experiment are to detect IDPs with masses between 10 16 g and 10 6 e

determination of IDPs spatial distribution, and determination of IDPs size distribution
and its spatial variation.

6.2.2.3 Model Payload for a Future Mission

The table below shows the model payload with a ten-percent margin from our present
knowledge about the instruments. r

Table 6.4 The Model Payload for a Future Mission.

Instrument

Mass (kg)

Power (W)

Telemetry

(kbits/s)

Typical Time
Resolution (s)

Magnetometer

4.0 (4.6)

3.0 (6.2)

4(4)

0.01

Plasma Waves

6.5 (6.7)

5.0 (5.1)

12 (12)

0.001

3 - D Ions

2.0 (3.1)

2.0 (2.7)

6(2)

0 . 1-1

3-D Electrons

2.0 (3.0)

2.0 (2.0)

4(4)

1

Heavy Ions

3.0 (3.5)

3.0 (3.5)

0.9 (0.1)

10

Superthermal Particles

2.0 (4.0)

25(2.5)

2.0 (0.7)

1

Energetic Particles

2.0 (3.5)

2.5 (4.0)

3.0 (2.0)

1 - 10

Dust

1.0 (1.2)

1.0 (1.0)

0.1 (0.1)

Total

22.5 (29.6)

21.0 (27.0)

32 (24.9)

—

6.2.3 Remote Sensing Instruments

6.2.3.1 Introduction

Optical imaging is a major tool for remotely studying the solar corona both from the
ground and from space and indeed coronographs have extensively contributed to its

understanding.

UV solar physics also has always been the centrepiece of solar research from space since
a) UV solar radiation smaller than 300 nm is absorbed by the Earth s atmosphere and
therefore can only be explored from space; b) the solar UV is the dominant energy source
of the upper Earth atmosphere; c) the outer solar atmosphere from the transition region
in jq the corona emits most of its radiation in the UV [Brueckner, 1993].

An ideal solar remote sensing instrument set has to combine:

• High angular resolution combined with the pointing stability.

• High spectral resolution.

• Good time resolution compatible with the angular resolution.

• Wide and simultaneous spectral coverage required to follow structures and
phenomena in velocity, density and temperature from chromosphere to
corona.

To meet all the scientific requirements involved in understanding the corona, a set of
instruments that complement each other and observe the same area or even the same fine
structure during co-ordinated programmes is needed. Their co-alignment will be more
and more difficult with the increasing angular resolution. Most of the instruments also
require high spacecraft attitude stability: upper limit of angular velocity for many of them
is of the order of 1"/sec.

Although it is not easy to combine high angular, spectral and time resolutions within the
physical constraints placed on space instruments, progress in this direction continues.
The most important trends in development of optical and UV devices are slow increase of
angular and spectral resolution along with steady decrease of mass and dimensions. The
latter parameters can not decrease as fast as those for electronic devices because they are
limited by optics, but there is a tendency to integrate several devices having different
bands (and even instruments from other spacecraft systems such as attitude
determination sensors).

To improve the quality of remote sensing measurements, it is necessary to improve a)
image quality and stability limited by quality of the optical surfaces and effects of
structural deformations; b) spectral resolution and stigmatism. Sufficient precautions are
to be taken to ensure that the image quality is optimised not only for ground tests, but
also for in-orbit constraints, by selecting the proper mechanical structure and mounts for
the optics and taking the spacecraft characteristics into account.

One more restriction is caused by negative effects of the solar environment on the
instruments and, especially, the optical surfaces. These effects can be reduced either by
placing the instruments behind tiny holes in the heat shield or by using retractable
mirrors which can be extended out of the umbra only for short measurement periods.

In this chapter only the instruments being designed for future solar probes, i.e. spacecraft
visiting the solar corona, are considered. Such an approach is caused by unique

requirements on such instruments and, hence, their significant difference from payloads
of spacecraft designed for 1 AU environment such as SOHO, Mir or Space Shuttle.

6.2.3.2 Remote Sensing Instruments from Russian Probe Plamya

Among the planned solar probe missions the Russian Plamya spacecraft has the most
advanced remote sensing instrument complex [Oraevsky, Kuznetsov, 19941. That is the
reason for considering all of them.

1 t P , ^ Se ,°! P ro P osed Plamya "Solar Coronograph” experiment is to construct a
global 3-D white light image (spectral range 5500...6000 A) as well as a global 3-D model
of the solar corona within 1.3...5 solar radii. The FOV of the instrument is 18°; angular
resolution is 2.1 per pixel. A series of white light coronal images recorded in various
projections during the Plamya passage between the two poles will be used to extract the
3-D structure of the entire solar corona. Using techniques similar to computer

omography a quantitative model of the solar corona can be derived from the Plamva
imaging data. y

According to [Vaisberg, 1996] Russian scientists and engineers managed to reduce the
mass of the instrument from 3 to 1.5 kg (without reducing capability).

The Solar Vector Magnetograph" experiment is designed to study solar magnetic fields
and radial velocities with spatial resolution of about 100 km within FOV of
100 000x100 000 km at a distance of 20 solar radii. On-board storage having 10 Mbits will
permit to obtain temporal resolution of 2-4 hours.

EUV Telescope with FOV 12° is modified version of the EUV channel (190-205 A) of RES-
C (solar X-ray spectrometer) operating on board of the CORONAS-I.

Plasma Analyser is to be mounted on a boom and has almost spherical field-of-view

6.2.3.3 Other Remote Sensing Instruments for Solar Probes

For future solar probes the Jet Propulsion Laboratory offers an instrument consisting of a
high resolution visible light telescope, a high resolution EUV telescope and two EUV
pinhole imagers combined in an integrated configuration [see table 6.5], The estimated
total cost of the instrument is 6.5 million US$.

Table 6.5 Characteristics of the JPL EUV/VIS Remote Sensing Instrument.

Band

Wave¬
length (A)

Spectral
resolution (A)

FOV

at 4 R s (km)

Spatial
resolution
at 4 R (km)

EUV

304

few A

100 000

390

304

few A

5 000

j 20

171

few A

100 000

390

Visible

4308

+10/-2

2 560

18

The Coronal Optical Imager is designed by Laboratorire d 'Astronomie Spatiale and
bishtu d Astrophysique, France [Lamy, Koutchmy, 1994], The instrument combines the
capability ot EUV, UV and visible imaging with spatial resolution of 100 km as well as
visib e polarimetry. The authors proposed two versions adapted to spinning and 3-axis
stabilised probes respectively. A Coronal Optical Imager (COI) can detect the faintest
plasma and magnetic structures, analyse the He/H ratio and the cool plasma component

Technology Challenges & Issues • 123

and observe possible sources of dust near the Sun. The instrument will have on-board
storage of 1.5 Gbits

The Solar Pioneer is a mission concept developed by Johns Hopkins University Applied
Physics Laboratory (JHU/APL) [McNutt et al, 1994]. Although the "core" set of
instruments for this mission includes only in situ measurements, there are also two
"strongly desired" remote sensing instruments.

Table 6.6 lists physical instrument characteristics for various programmes.

Table 6.6 Summary of Physical Characteristics of Remote Sensing Instruments.

Instrument

Mass (kg)

Power (W)

Data rate
(kbit/s)

Solar Pioneer Coronal Fhotometer

0.7

0.5

0.1

Solar Pioneer Coronal Disk Imager

3.3

5.5

0.4

Plamva White Light Solar Coronograph

3-»1.5

5

5

JPI intperated instrument (Vis. + UV)

4

1

0.35

mii i . ..

15

15

5

Plamya EUV Telescope

7

5

5

Plamva Vector-Magnetograph

7

8

5

Plamva Plasma Analyser

6

6

6.3 Orbit and Trajectory Definition

The objective of the orbit and trajectory analysis is to determine the flight profiles of
spacecraft subject to various constraints such as scientific orbital requirements, launch
time frames and minimisation of propellant expenditure. The Av budget (the sum of the
velocity changes required throughout the space mission life) is traditionally use to
account for the trajectory energy required throughout the whole mission. Various
trajectories will be discussed, and some elementary calculations will be made of the
velocities required for missions approaching the Sun. Due to the very high velocity
requirements, it will become apparent why until now, gravity assist trajectories have
been primarily selected. However, we will also explore the merits of low thrust
trajectories where propulsion is provided by electrically powered thrusters, solar sails, or
a combination. In addition, the possibilities of trajectory alteration due to aerodynamic
forces induced by planetary upper-atmospheric flight will be mentioned. This section
will provide a broad overview of solar-oriented trajectories, and will give some possible
trajectory options for the Ra missions.

6.3.1 Summary of Recommended Trajectories and Orbits for Ra

Alternative trajectory solutions were examined and the following conclusions and
recommendations were made. Further details are provided in the sections following this
summary. The emphasis of the analysis was on innovative solutions that did not rely
upon time extensive gravity assist and chemical propulsion manoeuvres that appear to be

the norm today.

Low thrust trajectories powered by solar sails or electric propulsion were examined as
well as the various combinations of gravity assist coupled with chemical or electrical
propulsion. Various orbits ranging from heliosynchronous to highly elliptical orbits out
of the ecliptic were considered.

i?4 • Ra- The Sun for Science and Humanity

The following is a list of conclusions drawn from the
own judgements and calculations.

reviews of the literature and of our

• Solar sails can provide a significant reduction in time of flight. Solar sail
trajectories have been analysed and should be considered as optimal solutions
for the long term plans, given the fact that many technological issues still need
to be addressed relating to the deployment and attitude control of extremely

u rS T e nf e rtf d , llShtWei8ht structures ’ B ^ed on some of the studies conducted
by JPL [Wright and Warmke, 1976], [Friedman et al, 1978] on solar sail
trajectories, it is feasible to reach highly inclined circular orbits at distances
within 0.3 AU of the Sun in time periods that are sometimes half of the required
time to make it through a Jupiter Gravity Assist (JGA). In addition, coplanar
(non-optimal) orbital transfers were examined with software we developed.
The key advantage of the solar sail lies in the fact that little or no propellant
mass nor power is required (compared to electrical and chemical propulsion) to
provide the necessary thrust. It is a fact, though, that attitude control of a
spacecraft equipped with large solar sails will require more attention and will
be much more complex.

• Interesting applications for Ra are presented in section 63.6.9 [Circular Orbit] and in
section 6.3.6.10 [Polar Eccentric Orbit].

• Electrical Propulsion (EP) provides low thrusts at very high specific impulses
and hence over a long period of time can deliver high velocity increments (high
Av). This technology is currently being applied to upcoming interplanetary
missions such as the New Millennium spacecraft developed by NASA and for
the Japanese Muses-C for its rendezvous with an asteroid. The electrical
propulsion should be seriously considered for the Mid-Term Programme.
Some solutions using electrical propulsion are proposed in the following
section of the report for the SAUNA mission [see chapter 9], since EP transfers
are efficient means to inject into high velocity circular orbits. EP provides a
reduced time of flight compared to gravity assist scenarios and a reduced
propellant mass compared to chemical propulsion. Depending upon the
thruster type and mission characteristics, the power requirements for thrust
production may be quite large. Solar array-powered EP thrusters are most
efficient when not too close to nor too far from the Sun and this has an impact
on the potential orbit selection.

• Interesting applications for Ra are given in section 63.6.6 [Electric Propulsion
Trajectories] and section 63.6.7 [Combinations of Direct Insertions].

• Gravity Assist (GA) flybys are conventional and low risk manoeuvres with a
proven historical heritage and therefore represent a viable solution for missions
in the short term. Given suitable planetary bodies, GA flyby manoeuvres can
provide a huge Av saving for the injection to elliptic heliocentric orbits.
However, very large or unfeasible Av's are required for the eventual orbit
circularisation. Compared to direct injection, solar sails or low thrust, GA
flybys introduce launch date constraints imposed by the required phasing of
the planetary bodies. Jupiter GA is the most effective flyby because any change
of inclination is attainable and very low perihelia can be achieved. However,
such trajectories imply a longer transfer time and can expose the spacecraft to
intense radiation environments. Venus and Mercury GA flybys, although less
effective due to the relatively lower mass of the planets, can provide sufficient
impulse to reach orbital inclinations up to 20 degrees. Nevertheless, perihelia

lower than the altitude of a Sun-synchronous orbit seem to be difficult to
achieve.

• An interesting example for Ra is given in section 6.3.6A [Highly Eccentric Orbit],
showing a resonant Venus flyby for a transfer to a 2 Ofi inclined orbit with a 0.25 AU
perihelion. GA transfers could possibly be integrated with electrical propulsion or solar
sail for the final orbit circularisation.

Preliminary Background and Information

The potential planetary bodies for gravity assist and their relative size and mass are given
in appendix C.5, table C.5.1. Table C.5.2 in appendix C.5 lists some of the distance units
that we will make extensive use of during this section of the report.

6.3.2 Orbit Review and Definition

Note that all the orbits evaluated in this chapter have the Sun as the principal focus of the
orbit. The most interesting possibilities for heliocentric orbits are studied and some of the
advantages of the various options are raised in this chapter.

6.3.2.1 Circular Sun Orbit in the Ecliptic (Eccentricity < 0.1)

These orbits are contained within the ecliptic and therefore can allow a study of the Sun
from low latitudes of the solar environment.

Sun-Svnchronous Orbit (0.18 AU orbit, ~28 days period)

These orbits are a subset of the circular Sun orbits, and allow the spacecraft to have a
period equal to the Sun's rotation (approximately 28 days, around the Equator and
increasing towards the Sun poles). This is achieved by sending the spacecraft into a
heliocentric orbit with a semi-major axis of approximately 0.18 AU from the Sun and a
relatively low eccentricity. Although this orbit appears to be quite close to the Sun
throughout the whole duration of its orbital path, it allows very interesting studies of the
solar environment, since the spacecraft can investigate the same point on the Sun by
turning around it with the same period and around itself once during one orbit.

Lagrangian Points (LI -> L5) and Halo Orbits

The Lagrangian points, or Libration points, for two celestial bodies in mutual revolution
are the five points such that an object placed at one of them will remain in essentially the
same position relative to the bodies. They are in the orbital plane of the two body system.
The motion about one of the stable Lagrangian point may be dominated by the
perturbation due to a third-body interactions. For the Sun-Earth system, LI, L2, L3 are
unstable points; that means if we place a body in one of these points, small correction
manoeuvres must be applied to prevent excessive departure from the nominal orbit. L4
and L5 are stable points.

L2 and L3 are of no interest for the Solar Probe mission since they are located in positions
where they either can not see the Sun or they cannot see the Earth (respectively).

Spacecraft (i.e. ISEE/ICE and SOHO) have been launched to orbit around LI, which is
located between the Sun and Earth at one hundredth of an AU from the Earth. Refer to
the various sections in this document and appendix C.5 to learn more about the trajectory
and orbit of SOHO. The spacecraft is actually orbiting around the LI point in a path that
we call a halo orbit.

. ? d L , 5 r c Very in ereshn § p0mts smce the y would eventually allow multiple view
points of the Sun and would provide images for building a fully integrated three-
dimensional (or at least fully two-dimensional) model of the Sun's environment.

6.3.2.2

Eccentric Sun Orbit (eccentricity > 0.1)

The main advantage of these orbits that are highly eccentric is that there is mainly no

rpT™^f C1 !£ U c nSe th f 0rblt l 0nce the spacecraft has reached the desired location with
respect to the Sun and once that it has been injected in the proper course. This implies of
course a much lower Av (in the order of 5-10 km/s instead of 25-35 km/s).

The main disadvantages lie in the fact that the period of the orbit is much longer, on the
order of 5 to 6 years depending on the aphelion, thus allowing close studies of the Sun
only during short periods when close to the perihelion.

6.3.2.3 Polar Orbit Around the Sun

This orbit has already been used by previous spacecraft and Ulysses is orbiting in a path
that takes it around the poles of the Sun in an eccentric orbit that brings it back to the
orbit of Jupiter where it had its trajectory modified through gravity assist. The main
interest of having a probe in a polar orbit around the Sun is that the polar environment of
the Sun is still quite unknown and would surely reveal a lot if we were to study the

presence and structure of solar magnetic fields and other solar events in the vicinity of
the solar poles. y

6.3.2.4 Heliocentric Geosynchronous (HGS) Orbit

This orbit is an heliocentric orbit with the orbital plane precessing at 1 deg/day to
maintain a fixed angle between the orbit plane and the Earth direction. The feasibility of
such orbits was initially investigated because of the obvious advantages that they would
offer for a prolonged mission. Earth Sun-synchronous orbits exploit the Earth oblateness
(term J 2 ) providing such an orbit plane precession for a given altitude and inclination
However HGS orbits are found not to be feasible because of the high sphericity of the

6.3.3 Achievable Orbits (Trajectory, Time, Energy)

The purpose of this section is to examine a broad range of trajectories to provide an
overview of how costly solar missions can be in terms of velocities required.

We will consider the possible following trajectories (a subset of which will be analysed in
higher detail in the following section): 1) direct injection; 2) gravity assisted (with or
without aerobraking); 3) low thrust with electric propulsion; 4) solar sail- and 5) a
combination of the mentioned orbits/techniques is also possible, e.g. gravity assist plus
low thrust or solar sail; solar sail plus low thrust etc.)

For more information on the various propulsion systems, please refer to section 6.4.

6.3.3.1 Gravity Assist and Aerobraking

An important consequence of a spacecraft entering a sphere of influence of a planet is the
possibility of gaining or losing energy with respect to the Sun (the vast majority of the
solar system s angular momentum is retained within the planets). It is this same
momentum that is used to accelerate spacecraft on so-called "gravity-assist" trajectories.

he gain or loss of energy is caused by the turning of the spacecraft velocity vector under
the influence of the gravitational field of the planet around which we perform the flyby.

The spacecraft's arrival date for the flyby needs to be carefully timed so that it would
pass close to the planet in its orbit around the Sun (optimisation software used for
trajectory definition are covered in a following sub-section). Gravity assists can be also
used to decelerate a spacecraft, by flying in front of a body in its orbit, transferring some
of the spacecraft's angular momentum to the body (negligible amount for the planet).
When the Galileo spacecraft arrived at Jupiter passing close in front of Io in its orbit,
Galileo experienced deceleration, helping it achieve Jupiter orbit insertion.

6.3.3.2 Chemical Propulsion (or Direct Injection)

Analytical formulas can be used for this purpose. The analysis is quite straightforward
and shows that with present technologies the huge Av's required for direct injection into
a Sun orbit (e.g. a highly elliptical Sun orbit or, more difficult, a circular Sun-stationary
orbit) is so costly to make this option not feasible (SOHO had only to be launched into
the Lagrangian point LI relatively close to the Earth rotating around the Sun with the
same period as the Earth). If a change of inclination is required (to go out of the ecliptic
plane to take high latitude measurements of the Sun), the situation is even worse. This
situation could however change in the future (though probably not in the short plan) if
new and more powerful launchers and upper stages are developed; therefore, this option
was not discarded during a first analysis and some results will be given in this report.

6.3.3.3 Ion Propulsion

Characteristics and advantages of ion propulsion are discussed in section 6.4.3.1

6.3.3.4 Solar Sails

Space sails use solar or other radiation directly as a method of propulsion. They are
large, lightweight mirrors which reflect either photons or electromagnetic radiation. The
advantage of using solar sails is that a power generator and converter are not necessary
onboard, thus saving mass and costs. The biggest disadvantage is the necessity of large
sails [refer to section 6.4.4.4J.

The trajectory course is determined by the departure and destination points, the
characteristic acceleration, the orientation of the sail, and by the thermal requirements of
the sail and the spacecraft.

Usually, the optimisation of interplanetary trajectories is based upon the minimisation of
the transfer time between the departure point and the final destination. This means the
sail angle must be optimised as a function of time. Orbital transfer optimisation with
solar sails has been studied at JPL, see for example [Sauer, 1976].

The initial conditions at the departure point depend on the speed and direction of motion
that the ship can have as it departs from planetary space into interplanetary space,
crossing the sphere of influence of the planet.

6.3.4 Orbits and Trajectories of Previous Solar Missions

For a review of the trajectories of the missions directed to the Sun that have been
accomplished, or are being conducted at the moment and that may even be planned for
the near future, refer to appendix C.5.

10 Q * Pd- Sun for Science and Humanity

6.3.5 Orbit Optimisation and Software Review

6.3.5.1 MIDAS

Please refer to the appendix C.3 on the MIDAS Software.

6.3.5.2 SKYNAV

Please refer to the appendix C.4 on the SKYNAV Software.

6.3.5.3 Solar Sailing Optimisation Software

Please refer to the appendix C.2 on the Solar Sailing Software (called Sailing) to get the

i oc£ P xV° de m , FORTRAN g ener aled during the Summer Session Program in Vienna
rake n°te that the code was based on a previous program written during the ISU
session of 1994 in Barcelona for the study of Mars Aerobraking and that it does not
perform any optimisation.

6.3.6 Ra: Appropriate Orbits and Trajectories

In this section of the report, the examples that were analysed using the available tools
mentioned above are presented. Depending on the technology available (solar sails, low
thrusters etc.) some of these sample trajectories could be considered for the mid-term or
for the long-term missions.

6.3.6.1 Circular Orbits in the Ecliptic Plane using Direct Injection Trajectories

The first type of trajectories to be studied were the direct injection trajectories. To get the
orders of magnitude for these types of trajectories, a study of the different velocity
increments for various ecliptic heliocentric orbits was made. Varying the radius of the
orbit, we computed the first Av required first to reach the required orbit, and then the
second Av to circularise the trajectory to the final orbit.

Table 6.7 Velocity Increments for Various Ecliptic Heliocentric Orbits.

r p (Solar radii)

Av, (km/s)

Av 2 (km/s)

5

23.6

77.7

10

21.1

52.8

39

13.4

21.1

50

11.6

16.9

100

6.1

7.4

Using MIDAS software (by Carl Sauer and Stacy Weinstein of JPL) for trajectory
optimisation, some single and multiple planetary swing-bys were analysed. In
computations in sections 6.3.6.2 through 6.3.6.5, we assumed always to start from a
circular LEO parking orbit at about 200 km altitude.

6.3.6.2 Polar Highly Eccentric Orbit using Jupiter Gravity Assist (JGA)

Jupiter is often used for gravity assists when going to the Sun and cranking the orbit in a
polar and highly eccentric orbit (like the American Solar Probe) or in order to stay in the
ecliptic plane and further circularising. Using MIDAS, optimal Jupiter Gravity Assist

Technology Challenges & T.qjqhpc •

(JGA) trajectories were studied to go to a highly eccentric orbit with a very low perihelion
(about 4 solar radii R,).

A polar (90 deg inclined) orbit with a perihelion of 0.018 AU (4 solar radii, like Solar
Probe) was studied: taking 13/12/2004 as a launch date, a transfer time of about 4 years
was found to be necessary, and the launcher has to provide a C 3 —103 (km/sec) (for a
Av=7.2km/sec). The distance at flyby from Jupiter is 16 Jupiter radii, which should not
be a too harmful radiation environment. This is represented in figure 6.1.

Jupiter

18/09/Z006

Figure 6.1 Jupiter Swing-by with 90BInclination Change.

If the constraint on the perihelion of the target orbit is relaxed to be 0.18 AU (about 30
solar radii, similar to a polar version of the SAUNA option), the C 3 required decreases to
91 (for a Av of 6.8 km/sec).

6.3.6.3 Highly Eccentric Orbit in the Ecliptic Plane using JGA

Similar cases as in the previous section (final perihelion of 0.019 AU and 0.18 AU) were
analysed, while staying in the ecliptic plane. The required transfer time remained the
same (4 years approximately), as well as the C 3 requirements. The JGA to a final
perihelion of 0.18 AU in the ecliptic plane could in principle be used as a transfer orbit for
the SAUNA mission [see chapter 9] but the Av required for the final circularization is
very large (29 Km/sec). An alternative method with finite thrust starting from Jupiter
was not considered here because of the difficulty of providing the required electrical
power (from solar panels) at such high distances from the Sun [see below for
circularization with low thrusts].

6.3.6.4 Orbit Slightly Inclined (20°) using Resonant Flybys around Venus

This case considered is a resonant Venus Gravity Assist (RVGA) to go from the Earth
orbit to a solar orbit with a perihelion of 0.25 AU, 20 deg inclined with respect to the
ecliptic plane. This represents a rather cheap option and we found that using a resonant
gravity assist of this type, starting from a LEO, a C 3 =15.3 (km/sec ) 2 is required from the
launcher (providing a Av of about 3.8 Km/sec). The onboard propulsion must provide an
additional Av of 4 Km/sec at the time of the second Venus flyby. As expected, a further
circularization would need a very high impulsive Av at perhelion (13 Km/sec). Finite
thrust options could be studied for this latter purpose, since it is shown below that
circularization in a spiral low thrust trajectory starting from a Venus orbit (though with
different conditions) can be effective.

This case is illustrated in the following figure 6 . 2 .

Figure 6.2 Resonant Venus Flyby.

6.3.6.5 Circular and Elliptical Orbits using Multiple Mercury and Venus Flybys

Such a technique was analysed in order to get to a Sun-synchronous orbit (defined as a
circular orbit at about 38 Solar Radii or 0.18 AU from the Sun, in the ecliptic plane). This
option allows a relatively fast transfer, but imposes very costly requirements both on
auncher and onboard propulsion. Selecting the launch date on 29 December 2000 the
aiuncher has to provide a C 3 =113 (km/sec ) 2 from a LEO (meaning a Av =7.5 Km/sec)
Mercury and Venus orbits will then be encountered respectively after 68 and 115 days
(both flybys being unpowered), and two additional manoeuvres of 2.5 and 19 Km/sec are
needed after the Mercury flyby in order to respectively get to the desired perhelion and
then circularise. The launcher requirements would make the launch very expensive
r ? qui . n Q ng P° s " ,b 'y a pr OTON launcher with upper stage). The final orbit is reached after
a out 8 months from launch. The use of low thrust for circularization after the Mercury
flyby was also considered and optimised (using the software SKYNAV) but did not eive
an interesting result (probably due to the eccentricity of the spacecraft orbit after the
Mercury flyby and to the vicinity of the Sun, see figure 6 . 3 ). About three more months
are needed after the last Mercury flyby, a thrust level of 1.3 N for a total Av of about 40
Km/sec. Even using 4 Plasma Xenon thrusters (providing 0.3 N each) with a relatively
lgh specific impulse (2000 sec), the mass and power requirements (60 kW for

propulsion) make this option not recommendable. This example is illustrated in the
following figure.

Earth 2Q / 2 2/2000

Mercury 23/04/200 2

Figure 6.3 Venus-Mercury Flyby followed by Electrical Propulsion.

Figure 6.4 Top View (Ref. Radius). Figure 6.5 Av Requirement (m/s).

We used the same technique to go to a distance of 30 R s (solar radii) or 0.14 AU from the
Sun (circular or not), changing the inclination to about 20 deg. The launch date was set to
29 December 2000. Starting as usual from a LEO, the launcher has to provide a C 3 -113
(km/sec) 2 and the two planetary flybys are unpowered. An additional Av of 5.6 Km/sec
lowers the perhelion to 30 R, and modifies the inclination, whereas a huge final Av = 24
Km/sec would be required for the final circularisation. The total transfer time is 8

months.

6.3.6.6 Electric Propulsion Trajectories

Using a special software (SKYNAV) available at ISU some low thrust trajectories were
considered. This technique can also be considered in conjunction with GA flybys or with
direct injection, since in these cases the eventual circularization with impulsive
manoeuvres is very expensive.

J P h acps to m T s j ; 0 " r in s^

1» HC p,ane ' This orbit ,s similar to ,hat ™

Direct injection into the orbit. The low thrust propulsion system was chosen to consist of
xenon ion thrusters with a combined thrust level of 0.28 N. The transfer time was about

final mass faction'offs? W> a " d ' he ‘ Sp ' S 6000 rasultl "8 wilh a «***

This trajectory is illustrated in figures 6.4 and 6.5.

Figure 6.6 Top View (Ref. Radius).

Figure 6.7 Av Requirement (m/s).

6.3.6.7

Combinations of Direct Injections and Electric Propulsion

The second low thrust trajectory uses an initial velocity increment to Venus' orbital

‘f,; Ce f "° m S -P r A ov ! ded *7 a Puncher and upper stage. Upon arrival at this orbit,
low thrust is activated A thrust level of 0.27 N and an Isp of 4500 s is assumed. The total
Av is the same as in the previous case, but the total transfer time is one year. This could

be possibly improved by incorporating a simple or a resonant gravity assist at Venus [as
in section 6.3.6.4J.

This trajectory is illustrated in figures 6.8 and 6.9.

Figure 6.8 Top View (Ref. Radius). Figure 6.9 Av Requirement (m/s).

Technology Challenges & Issues • 133

6.3.6.8 Lagrangian Point L4 and L5 using Conventional Direct Injection

The easiest way to make it to L4 using a conventional direct injection trajectory, would be
achieved by allowing the spacecraft to go from a parking LEO orbit to a heliocentric orbit
that would have a lower period than the Earth for some time and to then re-inject it back
to a 1 AU orbit. This would take 5/6 of a year (0.833 year) and would require
approximately 3.885 km/s of Av through a transfer to an orbit with a semi-major axis of
1.32x108 km. Please refer to figure 6.10 for more information and visualisation of the

scenario.

The same conceptual approach to reach the L5 would lead to allowing the spacecraft to
go to a higher orbit than the Earth until it actually reaches the desired location (by taking
some delay on the Earth during their orbit) and would then be re-injected to a 1 AU orbit.
The process would then take 7/6 years (1.167 years) and would require approximately
2.843 km/s of Av through a transfer to orbit with a semi-major axis of 1.65x108 km.

Figure 6.10 The Five Lagrangian Points.

6.3.6.9 Circular Orbit at 0.18 AU using Solar Sailing

Using the Solar Sailing Software (written at ISU), we have tried to model the nature of a
trajectory for a spacecraft leaving the Earth and spiralling towards the Sun using only
solar sailing as the main propulsion. We have assumed that the sails maintain a constant
angle of attack with respect to the incoming solar pressure and that the trajectory could
be contained in a two-dimensional plane.

To provide the report with an order of magnitude, we have computed the trajectory for a
spacecraft leaving a parking orbit, with a mass of 250 kg, and solar sails of 9000 meter
square (that represents 30 meters xl50 meters sails on both sides of the spacecraft). The
results of the calculations show that the spacecraft would require approximately 928 to
937 days (or approximately 2.55 years) to make it to a 0.18 AU orbit around the Sun if the
sails are maintained at a 45 degree angle of attack throughout the complete spiral. We
have also noticed that the inclination of the sail has a great impact on the nature of the
orbit If the sail was to be inclined with a 60 degree angle, than the spacecraft would
never manage to reach the 0.18 AU in 950 days, but if the sail was oriented with a 30
degree inclination, then the 0.18 AU orbit would be reached in less than 546 days
(approximately 1.49 years). Please refer to appendix C.2 for more information on the
computations and on the mathematical equations used.

We also repeated the orbit propagation with solar sail when starting from Venus circular
orbit and ending with the Sun-synchronous orbit: a total transfer time of 540 days is
found, considerably longer than with low thrusters [section 6.3.6.6].

6.3.6.10 Polar Eccentric Orbit 0.2-1.4 AU using Solar Sailing

The h-ajectory to make it to a heliocentric orbit with a semi-major axis of 0.3 AU has been
studied using square solar sails of 250 m 2 . The spacecraft would first spiral all the way
through 0.3 AU, then crank the orbit to a 90° inclined orbit to the ecliptic. The orbit's
ap e ion would be raised to 1.4 AU and the perihelion would be lowered to 0.2 AU
therefore providing a final period of 2.7 years for the orbit. The example was provided
by the Jet Propulsion Laboratory (JPL) and is illustrated in figure 6.11.

v

006 km 2 * 2

DLAr

27.9 ceg

AX*

1.25 mm/» 2

Figure 6.11 Polar Eccentric Orbit 0.2-1.4 AU using Solar Sailing.

6.4 Propulsion

An important subsystem of spacecraft is the propulsion system whose function may be to
provide any or all of the following: orbit insertion, orbit correction, attitude control or
even as in the case of electric propulsion or solar sails, to provide low thrust over an
extended period of time to achieve large velocity increments. As was discussed in section
6.3, the velocity increments required for orbits close to the Sun are very high. The
purpose of this section is to discuss briefly the various propulsion technologies that are or
will be available for use on proposed Ra missions in the various time frames and to make
recommendations concerning the most suitable propulsion systems.

6.4.1 Summary of Recommended Propulsion Systems for Ra

Chemical propulsion is extensively used in main and attitude control propulsion for
upperstages and spacecraft. What makes chemical propulsion attractive is its relatively
low amount of development testing and cost. However, chemical propulsion
performance is not as good as many advanced propulsion systems. If very high Av is
required and a planetary assist manoeuvre is not a viable option, then the use of chemical
systems for main propulsion could be prohibitive from a volume and mass stand-point

However, chemical propulsion, with its extensive flight heritage is the best option for the
attitude control system. r

Although electric propulsion (EP) does not have an established flight heritage, especially
for scientific missions, it is an attractive choice for missions that require large velocity
increments due to the low propellant consumption rates. Indeed, EP is an enabling
technology to reach inner solar orbits without the use of gravity assist manoeuvres. The
most viable current EP system that is available for lower power applications is an ion
thruster, and both U.S. (New Millennium) and Japanese (Muses-C) missions are planning
on using such systems in the intermediate future. In the intermediate and far term time
frames, it is strongly urged that ion thrusters be used as primary propulsion on
trajectories such as direct injection into circular heliocentric orbits < 1 AU.

Photonic propulsion is a very efficient and cost effective way for solar exploration
because 1) its performance increases with decreasing distance to the Sun; 2) it can easily
change the orbital plane when it is close to the Sun; 3) it needs no onboard power plants
to obtain propulsion effects; and 4) it does not consume propellant. However, solar sails
still require much development work, especially in the field of the ultra-light structural
design and deployment techniques. Therefore, this propulsion technology will be
available only for the far term program. As this program is supposed to go close to the
Sun, including suicide probes, it will able to exploit the advantages of photonic
propulsion. Together with nanotechnology which reduces the mass of the spacecraft and
also the area of the solar sail significantly, solar sails will represent a substantial leap in
technology for propulsion for future solar explorations. Below, the various propulsion
technologies are discussed in greater detail.

6.4.2 Chemical Propulsion

When selecting a propulsion system for a spacecraft, one option often considered is
chemical propulsion. This is especially true in today's environment where space budgets
are decreasing, and the use of existing technologies, such as chemical propulsion, is
encouraged and often required. Chemical propulsion has been the dominant propulsion
technology since the beginning of the space program. This extensive heritage and the
possibility of minimal development testing and cost, will continue to make chemical
propulsion an attractive choice. However, the heritage of a propulsion technology is not
enough to justify its selection, if its performance is not adequate to fulfil mission
requirements, or poor performance results in a heavy, large volume spacecraft and hence
large development costs. This chapter will provide information on various chemical
propulsion technologies including their advantages and disadvantages. A summary of
typical performance data for various chemical propulsion technologies is included [see
table 6.8]. A summary of advantages and disadvantages of various chemical propulsion
technologies is included in table 6.9. For a more detailed description of each propulsion
system, the reader is encouraged to refer to [Sutton, 1992] and [Wiley and Wertz, 1992].

Table 6 8 Typical Performance Data for Various Chemical Propulsion Technologies
[Larson and Wertz, 1992], [Price, 19961.

Propellant

Thrust Range (N)

Average Bulk Density
(ll/cm 1 )

Vacuum I ip (sec)

LO./LH,

5 - 5 x 10 6

1.14/0.07

450

GO./GH,

110-890

1.14/0.07**

440

5 - 5 x 10 s

1.43/0.086

300-340

0.05-0.5

1.0

150-225

Cold Gas (GN,)

0.05 - 200

0.28*

50-75

* At 24 MPa and 0°C. ** Stored as a liquid

Table 6.9 Advantages and Disadvantages of Various Chemical Propulsion
Technologies [Price, 1996].

Propellant

Advantages

Disadvantages

lo 2 /lh 2

• Extensive Heritage

• Best Performance of Chemical
Options

• Complex Engines

• Thermal Challenge of Keeping
Propellants at Liquid Phase

• Large Propellant Storage Volumes
(Especially LHd

GOj/GHj

• Good Performance

• Potential For Integration With a
L0 2 /LH 2 System

• No Flight Heritage-Potentially

High Development Cost

• Challenge of How to Best Get
GO,/GH, From LCL/LH,

NTO/MMH:

• Extensive Heritage

• Best Performance of the Storable
Technologies

• No Ignition System

• Corrosive Propellants

• Risk of Upstream Ignition of MMH
and NTO Propellant Vapours

• Compatible With a Limited Range
of Materials

N 2 H 4

• Extensive Heritage

• No Ignition System

• One Propellant-No Risk of
Upstream Ignition

• Positive Expulsion Propellant
Acquisition

• Corrosive Propellant

• Ammonia Dissociation Removes
Energy From Exhaust Gas

• Compatible With a Limited Range
of Materials

Cold Gas
(CN 2 )

• Good Heritage

• Lowest Performance of the
Chemical Propulsion

Technologies

• Non-toxic

• Simple, No Combustion

• Large Volume Required For
Propellant Storage

• Performance Degrades With Time

For Blow Down Systems.

6.4.3 Electric Propulsion Systems

Electric propulsion (EP) is increasingly being considered for a variety of applications
ranging from technology demonstrators to science missions and commercial applications
such as station-keeping on geostationary communications satellites. EP is attractive
because the advantage of the high exhaust velocities is significant on reducing propellant

However, in addition to increased propulsion system performance, spacecraft designers
and integrators must also consider the unique and important issues of spacecraft
contamination by EP thruster plume backflow, and how EP thrusters modify the
environment surrounding the spacecraft. This aspect is particularly important for
scientific spacecraft that must take sensitive measurements. If these issues prove
problemahc, it will be possible to cycle thruster operation and scientific measurements.

EP thrusters have traditionally been divided into electrostatic (ion and Hall)
electromagnetic (magnetoplasmadynamic), and electrothermal (arciet and resistojet)
types. Based upon evaluations considering the specific impulses, efficiencies and power
levels required, ion and Hall thrusters are the preferred thrusters for high velocitv
increment missions. They also are relatively advanced in their state of technology

readiness and have been selected for future missions (NASA New Millennium and
Japan s Muses-C).

6.4.3.1 Ion Thrusters

In ion thrusters, ions are formed in a chamber either by electron bombardment radio
requency excitation, or surface contact ionisation. These ions are then extracted and
accelerated as a beam to very high velocities (>10 km/s) by a system of highly charged

Tprhnrrinrri/ e t

grids. To maintain charge neutrality (and current balance for the spacecraft), electrons
are injected into the beam.

An example of a current ion thruster is the 30 cm diameter NASA Solar Electric
Propulsion Technology Applications Readiness (NSTAR) ion thruster that utilises xenon
propellant. The thruster is throttleable, and operating conditions can range from power
levels of 0.7-4.9 kW, with thrusts from 28 to 178 mN, and exhaust velocities ranging from
28-45 km/s. This thruster is currently planned for use on upcoming NASA New
Millennium missions, with the performance constrained to a peak power of 2.3 kW
(throttle range 0.49-2.3 kW), a thrust level of 20-92 mN, and a specific impulse of 2000-
3300 s. The NSTAR thruster has a mass of 8 kg (power processing unit is an additional 13
kg), an efficiency of 92% at full power, and a lifetime of 8000 hours at full power.

NASA has also developed a 14 cm diameter, 1 kg (power processing unit is an additional
8.5 kg) ultra-light NSTAR derivative that operates at a peak power of 1 kW (throttle range
0.25-1 kW), a thrust level of 30 mN, and a specific impulse of 3300 s. At full power, the
efficiency is 85-90%, and the lifetime is 15,000 hours. In addition, the UK has developed a
number of ion thrusters which are described in other mission sections.

6.4.3.2 Hall Thrusters

Hall or Stationary Plasma Thrusters (SPT) are also attractive propulsive systems for lower
power, high specific impulse missions. SPT's are essentially gridless ion thrusters that
make use of the jxB force. Propellant, typically xenon, is fed between two concentric
cylinders in which a gas discharge takes place. Magnetic coils create a nearly radial
magnetic field on the order of 150-200 G. An axial electric field is applied, on the order of
100-700 V, which generates an azimuthal Hall current in the ExB direction. This current
interacts with the magnetic field, producing a volumetric jxB accelerating force on the
plasma. Since the magnetic field is sufficiently weak that the ion gyroradius is much
larger than the dimensions of the thruster, the ions are accelerated to nearly the full
applied potential. The absence of grids, and a quasi-neutral plasma means that the
current-limited condition of conventional ion thrusters is not experienced. Similar to ion
thrusters, the plumes of SPT's contain fast beam ions, neutral propellant, slow CEX ions,
and sputtered electrode material. Xenon is the most common propellant, and between
50-100 SPT's have been used onboard Russian spacecraft over the last twenty years for
attitude control [Wetch et al, 1991].

Currently, the U.S. Ballistic Missile Defence Organisation (BMDO) and NASA are
proposing a mission for a solar electric powered spacecraft that will test a Russian Hall
Thruster with a specific impulse of 1600 s, and an efficiency of 50%. Other versions of
this thruster, sometimes referred to as a thruster with anode layer (TAL) have a peak
power of 10 kW (throttle range 2-10 kW), a thrust level of 0.25 N with a specific impulse
of 2500-3000 s. Thruster mass is 8 kg, with a lifetime of 5000 hours.

6.4.4 Solar Sails (Photonic Propulsion)

Besides chemical and electric propulsion, photonic propulsion using solar sails is a major
challenge for propulsion techniques used in solar missions. Photonic propulsion is a
unique technique because it uses the Sun as the major energy source which, unlike rocket
fuel, is free and unlimited. The use of solar sails is most effective for missions in the inner
solar system due to the increase of solar flux with an decreasing distance from the Sun.
Therefore, it represents one of the most attractive propulsion systems for solar missions

close to the Sun.

138 • Ra: The Sun for Science and Humanity

Fundamentally, the concept behind the solar sail [Wright, 1992] is to use large reflective
surfaces to provide propulsion for the spacecraft through the use of Sunlight pressure
(solar photon flux) for the motive force. This force F is generated by the process of
collision and reflectance of photons with the reflecting surface of the sail, see figure 6 12

and can be roughly approximated by using the equation

F = AQ tun (\ +q )^^l (6.1)

c

where A is the total area of the sail, Q Sun is the solar photon flux [see section 6.7], q is the
reflectance of the sail, <p is the angle of incidence of the solar sail [see figure 6 12] and c is
the velocity of light.

Figure 6.12 Schematic Representation of a Solar Sail with a Tilting Angle for
Decreasing the Velocity of the Spacecraft.

In order to reach locations in the vicinity of the Sun, sunlight must be reflected ahead of
the sail along its orbital path. This creates a negative force component along the sail's
path which pushes the spacecraft back and reduces its velocity [Diedrich, 1996]. In
addition to manoeuvring in the ecliptic plane, solar sails can also change the orbital plane
around the Sun easily by turning the sail so that the lateral acceleration is out of the
of the Su 3ne 11118 ° rblt Changing ca P abilit y allows the investigation of the polar regions

The major advantage of solar sails used as a propulsion system in solar missions is that
mass and cost of the spacecraft can be reduced significantly due to the absence of an
onboard generated power system which provides the propulsive effect [Friedman, 1988].
ITowever with solar sailing the change in velocity is applied very slowly but constantly
which leads to long mission times for achieving large values of Av. For a reasonable
acceleration by using photonic propulsion for solar missions, large solar sails must be
considered which span over an area of several square kilometres depending on the
spacecraft's mass. This leads to technology challenges [Boisard, 1995] in many other
technical disciplines besides propulsion, e.g.

• Ultra-light structural design and analysis of large scale structures

• Material engineering and manufacturing for the sail and the supporting
structure

• Packaging and unfolding techniques for launch and deployment of the sail

• Reducing the payloads mass by using nanotechnology.

With today's technologies the use of solar sails is efficient in a temperature range roughly
(because it depends on the material) between -270°C and 400°C and the minimum
approach distance to the Sun is about 0.06 AU [Wright, 1992], However, solar sails still
require much development work, especially in the field of the ultra-light structural
design and deployment techniques. Together with nanotechnology which reduces the
mass of the spacecraft and also the area of the solar sail significantly, solar sails will
represent a promising new propulsion technology that should be considered in far term

Ra programs.

6.4.5 Other Advanced Propulsion Concepts

An interesting propulsion technology that is being explored currently is solar thermal
propulsion, where solar energy is used to heat a working fluid that is expanded out in a
nozzle [Frye and Law, 1996]. Typical specific impulses range from 700-800 s. The term
"bimodal" systems is also commonly used to denote the use of a system that both
provides propulsion, and power via thermoelectric converters. Research on solar thermal
propulsion is being conducted at the U.S. Air Force's Phillips Laboratory at Edwards Air
Force Base. Development of a technology base for unconventional rocket thrusters using
intensely concentrated solar energy is currently in the exploratory development phase.

6.5 Power Systems

This section provides a brief discussion on power systems [section 6.5.1] and energy
storage [section 6.5.2] with respect to possible solar missions. Critical parameters to
consider for power system selection include the planned trajectory, the average electrical
power requirement, the peak electrical power requirement and the planned mission

life.

6.5.1 Power Sources

6.5.1.1 Photovoltaic Solar Arrays

A solar array is a very convenient method of converting energy and generating power. It
uses sunlight to convert energy directly into electricity. Important factors when selecting
solar arrays are temperature and degradation. Solar arrays are designed to work in
specific temperature ranges. The bonding between the arrays and the structure is also
temperature dependent and potentially critical. Solar array degradation is caused by
thermal cycling, micrometeoroid strikes, plume impingement from thrusters and material
outgassing. This depends heavily on radiation and is an important consideration near
the Sun. The fact that the solar flux varies with the distance from the Sun must be also

considered.

Efficiency

The efficiency is limited due to losses produced by sunlight reflection, conversion of
absorbed energy into heat and photon absorption. A comparison of the efficiencies of the
common solar array materials can be seen in table 6.10.

n _ * 'Tl— Ci

Qrionrp anH T-Tumanitv

Table 6,10 Solar Array Parameters [SMAD, 1992J.

Cell Type

Achieved

Efficiency

Degradation caused
by radiation

Problems

Silicon

14%

High

Low resistance to
high temperatures

Ga-As

18%

Medium

More mass, costs

Indium-Phosphate

19%

Low

Very high costs

There are basic methods to increase the efficiency to include, increasing the solar flux by
use of a concentrator [Scarlet Program, WWW] or by using a multi-layer or matrix design

6.5.1.2 Radioisotope Thermoelectric Generator (RTG)

A radioisotope (e.g. plutonium element) is used as a heat source [RTG Program, 1991],
Electricity is produced by a temperature gradient conversion method. Thermoelectric
coupling is a method of producing temperature gradients across materials of different
thermoelectric potentials in order to produce current flow [see section 6.5.1.5],
Thermionic energy conversion is a means of producing electricity by forcing an ionised
?n S t0 fon£° ra ^ e u, nC l COndenSe thrOU § h an eIectrical load. Typically efficiency is between
RTGs 20/0 ^ 6 6 11 Sh ° WS 3 feW ° f the advanta g es and disadvantages inherent in

Table 6.11 RTG Advantages and Disadvantages.

|_ Advantages _

Do not depend on
environmental or orbital
parameters

Disadvantages _

Politically very hard to handle
High specific costs
Radiation (requirements and
constraints about instruments)

6.5.1.3 Solar Thermal Dynamic Power Generation

Never before has the thermal control of a spacecraft been combined with turbine power

generation in practice [French, 1996]. Here the technological possibilities and challenges
are discussed. 5

A closed steam cycle is used where a part of the excess heat is transformed into electrical
power [figure 6.13]. The heat from the instrument or spacecraft system that needs to be
cooled is transferred to a fluid in evaporator tubes. (The tubes do not pass by the heat
shield or the multi-layer insulation because the small vessels would not support a
temperature higher than about 400 K and the correspondingly high pressure.) By the
fluid tubes the steam is transported to a turbine which expands the steam. No steam
should condense in the turbine because this would cause blade damage. Condensation
takes place in a radiator afterwards. The turbine does not only drive a generator but also
a pump that transports the liquid back to the heat source. The turbine will be switched
on only at a certain distance from the Sun, when there is the coincidence of the

requirements for excess heat to deal with and instruments needing electricity to do
measurements. y

evaporator tubes

Figure 6.13 Closed Steam Cycle.

We estimate that a single-stage, single-valve turbine, for instance, though it has a lower
efficiency (about r|=0.3) than multistage turbines could generate 15 kW of power with a
mass flow of 0.5 kg/s. So, the turbine could provide the power for the small set of
instruments for a spacecraft. Some investigations about a turbine cycle were done in the
Solar System Exploration [ISU, 1994]. The turbine/generation system described consisted
of copper straps and aluminium heat exchangers and tubes. The system was reliable for
one year and had a 20 - 30 kg mass.

6.5.1.4 Stirling Engine

Figure 6.14 Stirling Engine Principle.

The Stirling engine fulfils power and thermal protection needs. It can be applied as a
heat engine cycle, in which heat is accepted at a high temperature, rejected at a lower
temperature, and work or power is produced. A mirror collects the heat and transfers it
to the Stirling cycle. The Stirling cycle engine has proved to be the most promising
candidate of various solar thermodynamic cycles [Egushi, 1990] because its efficiency
(around 27%) is much higher compared to for example Peltier elements (5%). Figure 6.15
shows the Stirling cycle, the useful work produced by the cycle is represented by the
areas inside the P-V and T-S diagrams.

Figure 6.15 P-V- and T-S Diagram of the Ideal Stirling Cycle.

Two space applications of Stirling engines have been proposed [Scarlet Program, WWW1.
One are the Small Radioisotope Stirling Engines, as the Stirling engine can be combined
wtth isotope power systems. The Stirling converter is able to achieve higher efficiencies
at these lower power levels, so a lower amount of isotope is required. The second

25 kWeat 9 So/ e [; ar f ^e-Piston Stirling Engines. Their power output is about

2 f 25 / ° OVera11 ef f lCienCy - Heater tem peratures of 1050 K and cooler temperatures

of 525 K have been tested. Lifetime was 60000 hours, specific mass 6 kg/kWe.

So, the Stirling engine would be an effective power and thermal control device for a solar
mission The power output is higher than our first estimate of the fluid turbine's. Heater
and cooler temperature are not limited like in the turbine cycle (400 K temperature limit
of the flmd tubes). The heat-collecting mirror can be mounted just behind the heat shield

? r 8 /uw^ PeratU u e ,S radi ent can be reached. But the fluid turbine's specific mass (1.3-
2 kg/kWe) is much lower than the surveyed Stirling engine's (6 kg/kWe).

6.5.1.5 Peltier Elements Power Generation

cofd side

Figure 6.16 Peltier Element.

Peltier elements are a means of thermoelectric coupling [see figure 6.16]. Advantages are
the very rapid heating and cooling and the precise temperature control, as well as
simplicity and reliability. No moving parts and refrigerants are required, which means
there is no mechanical wear out, no danger of mechanical damage and less mass
compared to the Stirling engine or the fluid turbine. But, the temperature range would
also be different. While the Stirling engine is applicable everywhere behind the heat
shield and the turbine between 400 K and the instrument temperature, Peltier elements
operate at room temperature to -100°C. That qualifies them as a follow-up device for
precise instrument cooling, maybe for infrared and gamma ray detectors. More
widespread use of Peltier elements is limited by their very low efficiency.

6.5.1.6 Electrodynamic Tethers

This method uses the magnetic field of planets or the Sun [figure 6.17] to produce
electricity. The spacecraft has to provide a large, thin and isolated wire which crosses the
magnetic field. This induces a current in the wire and therefore power. Additionally to
the isolated wire an electron gun must be used to make the current constantly flow.
Limitations include: high voltage (20kV-40kV) isolation at the tether and the spacecraft;
high power, high-to-low voltage converter; plasma-electrodynamic interactions affecting
return current losses; the current produces a small thrust which must be added to the
trajectory calculations; and all past attempts including those of recent missions have
failed due to defects in the deployment mechanism and/or deployment.

Figure 6.17 Electrodynamic Tether [Tethers in Space, 1983].

6.5.1.7 Comparison

Table 6.12 shows a comparison of the discussed power systems. Figures 6.18 and 6.19
relate specific power and specific cost, respectively.

Table 6.12 Power Source Comparison.

Power Source

Efficiency

Max.

Temp.

Boundary

Conditions

Qualification Status

Solar Arrays
Silicon

14%

700 K

Spaceflight approved

Ga-As

18%

780 K

In-Phosphate

19%

680 K

RTG

7%

heat shield

Spaceflight approved

Fluid Turbine

30%

400 K

Fluid Tubes

Not yet qualified

Stirling

25%

1050 K

Mirror

Not spaceflight approved

Peltier Elements

7%

300 K

High Temperature
Gradient

Spaceflight approved

Electrodynamic

Tethers

99%

1000 K

Needs strong
magnetic field

liiK

1/1/1 *

t?-.. Qnn fr»r ^ripncp and Humanitv

Solar Solar RTG Stirlin
Photo Therm g

20000
15000
$/kg 10000
5000
0

Solar Solar RTG
Photovo Thermal

Figure 6.18 Specific Power [SMAD, 1992],

Figure 6.19 Specific Costs.

6.5.2 Energy Storage

Energy storage is an integral part of the spacecraft's electrical-power subsystem. Any
spacecraft that uses photovoltaics as a power source requires a system to store energy for
peak-power demands and eclipse periods. Primary batteries [table 6.13] are mainly used
for memory backup systems, which use very little power. They convert chemical energv
into electrical energy but cannot reverse this conversion, so they cannot be recharged.
Secondary batteries [table 6.14] convert chemical energy into electrical energy during
discharge and electrical energy into chemical energy during charge. An important factor
is the depth-of-discharge (DOD). It is the percent of total battery capacity removed
during a discharge period. Higher percentages imply shorter cycle life. Finally, fuel cells
store energy by water electrolysis. Fuel cells combine the two gases again and produce
electricity. The main advantage is, that it is not significantly influenced by the power
flux. One the other hand it has much more mass and is more difficult to operate.

Table 6.13 Primary Battery [SMAD, 1992].

Primary Battery Couple

Specific Energy
Density (W hr/kg)

Typical Application

Silver zinc

60-130

High rate, short life

Lithium thionyl chloride

175-440

Medium rate, moderate life

Lithium sulphur dioxide

130-350

Low/medium rate, lone life

Lithium monoflouride

130-350

Low rate, lone life

Thermal

90-200

High rate, very short life

Table 6.14 Secondary Battery [Technology for Small Spacecraft, 1994].

Secondary Battery
Type

Specific Energy
Density [W hr/kg]

Lifetime Cycles
(at 50% DOD)

Qualification Status

Nickel-Hydrogen

29

5000

Spaceflight approved

Lithium-Carbon

60

1200

Spaceflight qualified

Lithium-Ion

90

r 1000

Not Spaceflight approved

6.6 Structures and Materials

During the solar mission the spacecraft will meet very powerful multiform influence of
solar environment like huge heat, gases flows and radiation. In such conditions one of the
main questions is how to protect delicate instruments and at the same time give them full
ability to provide all necessary measurements. To reach these goals we have to pay great
attention to the new advanced materials with unique properties and to the structure

which have to provide stable conditions inside the spacecraft during the entire solar
mission.

6.6.1 Structures

The goal of this chapter is to show all the requirements on the structure and to describe
two main varieties of structures that can be used in this case, including their advantages
and disadvantages.

6.6.1.1 Requirements

The structures used for our solar missions must meet many requirements. The
requirements vary with the proposed missions. These requirements are stringent because
the solar environment is very harsh and because there is very sensitive equipment on the
spacecraft. We can distribute the requirements into several groups as shown in following

table:

Table 6.15 General Requirements.

Caused by

Requirement

Description

Environment

Heat protection

Both equipment and structure must be protected.
The deformations due to extreme heat can be very
complicated and unpredictable.

To avoid high stress concentrations, the
temperature should be distributed equally among
all the elements of the structure.

Gases protection

Physical properties of the structure can be changed
under action of hot gases and possible chemical
reactions. If we cannot ensure that changes are not
dangerous the only way is providing necessary
protection

Radiation protection

The action of radiation on the structure is similar
to gases action but it's more predictable and
depends on working time.

Equipment

Certain stiffness

The particle's flows of different density can cause
dangerous oscillations and disturb work of
systems. To avoid that we have to use right
materials and also provide necessary geometry of
the construction able to damper the vibrations.

Certain strength

We are not able to avoid absolutely temperature
gradient and internal forces in the structure. So
this requirement serves to minimise deformations.

Certain stability

We have to remember about long elements of the
structure and provide the necessary cross section
area because under action of deformations they
can lose their form

Ability to provide
accurate measurements

The structure has to have additional mechanisms
to provide especial conditions during the short
periods of time for extremely accurate devices

Certain alignments of
devices and distances
between them

Close location of different devices can cause a
conflict among them. Their needs are very special
and the structure is to provide necessary
conditions for each system

Launch vehicle

Minimum mass

Apart from using advanced light materials we can
compute the optimum geometry of the structure to
make whole mass as less as possible

Accessible geometry

Under the payload fairing we'll locate special
damper system to avoid strong actions during the
launch. So sizes of the payload are very limited

Ability to stand
overloads

Because of complicated trajectory we have to turn
on the engines several times. So the structure has
to stand all shifts of external forces.

Technology

Possibility to produce

Implementation of new materials and engineering
solutions will force to find new technologies and
check their reliability

Managers

Cost

The work on solar mission supposes to provide a
lot of researches and tests those can be expensive

ma • Ra-The Sun for Science and Humanity

6.6.1.2 Two Possible Types of Structures

Frame

Unified Volume

Figure 6.20 Frame and Unified Volume Spacecraft Structure.

Two quite different structure's types are examined: Frame and Unified Volume. Both
types are illustrated in figure 6.20. In case of Frame all instruments are independent
enough and the only connection the other ones is the frame construction itself. Using the
Unified Volume we have everything in one protected box. These types have their
advantages and disadvantages represented in tables 6.16 and 6.17:

Table 6.16 Frame,

Advantages

1 Disadvantages

El W!f

We can relatively avoid useless and harmful
interactions among the sensitive devices and so
simulate only natural environment surrounding
the sensor

vAIl necessary protection is individual

Under action of gases flow the structure is
getting disturbances. To avoid vibrations we
must use a lot of additional bars increasing the
number of connections among the main frames.

It makes the construction heavier.

• No general protection

Each device may require its own level of
protection and access to the environment.

In this case we also can reach minimal mass of
the whole spacecraft.

• Small mass

Many elements of the structure may require
similar kind of protection but it is unique for
each element and can be sometimes
implemented with difficulties because of
complicated shapes and big summary surface
area of the elements.

Usually frame type of the structure provides
minimal weight

Table 6.17 Unified Volume.

Advantages
L Small ine rtia moment

Disadvantages

iHflimful influences on the instruments

To provide orientation control in conditions of strong
solar influence is very important task. If onboard power
amount is limited the SC has to spend as less as possible
for control purposes. So inertia moment will determine
the minimal expenditure of energy for orientation and
stabilisation
—Tasv production

Unified volume means also unified standard protection
covering the whole surface and having maybe several
complicated openings for systems' tasks.

• Low cost

Production of such SC type for solar mission is more
traditional and can use wide experience._

Reflection of external radiation inside the box, not
perfect sealing from heat gases, interactions among the
devices and systems located too closely because of
limited volume can cause errors in their work

•. Temperature deformations

In solar environment strong heat at the one side and no
heat at the other one causes large temperature gradients
which cannot be absolutely reduced by
thermoregulation system. So hot surfaces of relatively
big sizes can lead the structure to significant
deformations which could be avoided by additional
heavy elements.

6.6.2. Materials

For the survival and proper function of a spacecraft used for a solar mission applied
materials and structure must meet extreme requirements for adequate resistance to the
harsh space environment. This chapter will provide a description of this environment as
well as the requirements. In addition an evaluation of materials which are suitable for
different kinds of solar missions is included in this section.

6.6.2.I. Environmental Conditions

The effect of the mission environment on structures must be considered in terms of both
the role of the structure and its operational life requirements. The environments to be
considered when selecting the appropriate materials are.

• manufacturing, transport, and storage

• launch

• in orbit space environment

We will only focus on the orbital environment of a solar mission, as the first two items do
not differ from others. However, they will be included for the material evaluation. The
material selection must consider the following:

• vacuum: primary concerns are the sublimation of metals, outgassing,
offgassing, and surface contamination in deep space

• radiation: particle and ultraviolet radiation, which becomes more severe
closer to the Sun

• temperature excursions, thermal cycling effects: the temperatures in a solar
mission can vary from -100 °C for some detectors to 2000 °C for a heat
shield at 4 R s . Changes in temperature influence both the mechanical and
physical properties of materials.

• space debris: assumed density of 2.8 g/cm 3 . It must only be considered in
the very early phase of a solar mission.

• micrometeoroids: average density of 0.5 g/cm 3 ' at altitudes higher than
1000 km. Increased dust densities close to the Sun. Although the average
density is much lower than that of space debris, impacts of micrometeoroids
must be considered during the whole solar mission.

These effects act together, their intensity varying over the spacecraft surface. Their
negative effects on the materials' performance must be evaluated and counter measured,
especially for long term missions. Moreover, the scientific measurements of a solar
spacecraft can also be disturbed by:

• electromagnetic disturbances

• surface charging

• deep dielectric charging

• magnetic induction

6.6.2.2 Material Requirements for Spacecraft:

The selection of materials is basically governed by functions to be performed,
environmental factors and costs. Besides general requirements, in a solar mission, the
following requirements must be met:

• high specific strength

• high specific stiffness

high stability (resistance to buckling, cracking, corrosion, thermal loads)

• low thermal expansion

• appropriate thermal and electrical conductivity

• low outgassing

• high resistance against the space environment close to the Sun

For scientific spacecraft the selected materials must meet additional
ensure minimal disturbances of the measurements:

requirements to

• magnetic cleanliness

• electromagnetic cleanliness

• control of contaminants

6.6.2.3. Material Evaluation for Solar Missions

For solar scientific missions, which tend to utilise applied materials in a way which is at
the limit of the state of the art", the material selection and the development of new
materials is a major challenge. Materials with high specific properties (e.g. ratio of
strength or stiffness versus density) are attractive for creating mass efficient structures.
The primary choice of materials for present and future space structures is between light
metal alloy and polymer fibre composite materials. In high temperature or other hostile
environments other metallic, ceramic or specialised composite materials are appropriate
Figure 6.2 L shows a comparison between the specific properties of typical aerospace
materials. Steels are included for comparative purposes.

Figure 6.21: Specific Properties of Typical Aerospace Materials [Stonier, 1991J.

Alloys

Aluminium is relatively light in weight, strong, easily available, easy to machine and low
m raw material costs. In spite of their higher specific strength and stiffness, magnesium
and beryllium are difficult to machine. If harder structural materials are required, steel

Technology Challenges & Issues • 149

and titanium are selected. A major problem of light metal alloys is stress corrosion
cracking. Light metal alloys are applicable between 0 K and 1000 K. In this temperature
range, the sublimation of metals is not a major problem. So light metal alloys are a good
choice for secondary structure in any solar mission. Table 6.18 shows some material
characteristics of metal alloys.

Table 6.18: Material Characteristics of Light Metal Alloys [Turner, 19901.

Material

Density

Specific Stiffness

Specific Ultimate
Strength

Young's Modulus

Coefficient of
Thermal
Expansion

106(N*m/kg)

103 (N*m/kg)

{10-6/K)

Aluminium

2800

25

98.6

68000

22.5

Beryllium

1854

152

103.5

304000

11.5

Magnesium

1770

26

129.4

45000

25.2

Titanium

4428

25

187.5

110000

9

Composite Materials

Composite materials are a good choice for the primary structure of solar spacecraft
because of their unique combination of high specific strength and stiffness, good
dimensional stability and damping capacity, and low weight. Moreover, due to their
coefficient of thermal expansion which is close to zero they are best suited for high
precision structures. The general advantage of these materials is that the designer can
tailor and optimise the structure with respect to lightweight, strength, stiffness, and
temperature range by specifying 1.) the combination of the fibre and the matrix material
2.) the fibre volume fraction 3.) the number of plies and 4.) the fibre orientation angle of
the plies which constitute the laminate.

Fibre materials mainly used for spacecraft are carbon and Kevlar fibres. Carbon, boron,
silicon carbide, aluminium oxide, ceramics discontinuous fibres, whiskers, and particles
are used in metal matrix composites. Typical matrix materials are epoxy resins,
thermoplastics and metals.

Thermal Coatings

Spacecraft temperatures are strongly influenced by surface absorptivity and emissivity
values To reach the desired values in the solar environment, several types of surface
finishes can be used: Black paint coatings have a high absorptivity (=0.95) and emissivity
(—0 88) and are used to maximise the heat exchange between a surface and its
environment by radiation. White paints, in contrast, have a lower absorptivity (=0.25)
and high emissivity (=0.90). There are also other paints, film and tape coatings and metal
conversion coatings with very different properties available. Vapour deposited coatings
can reach very low absorptivities (=0.04) and emissivities (=0.03).

6.6.2.4. Hot Structures

For the SAUNA mission, which is expected to go to 0.2 AU distance from the Sun, and
the Suicide Probe which will go as close to the Sun as possible, until it is destroyed, hot
structural materials must be used. The development of these kinds of materials is a major
technological challenge.

i sn • Ra: The Sun for Science and Humanity

rigure 6.22

pecific Strength Versus Temperature for Metal- and Ceramic-Matrix
Composites. CFRP, carbon fibre reinforced polymers- GMC class matrix
composites; GCMC, elass-ceramic matrix comS
matrix composites; MMC, metal-matrix composites* C-C carbon rarhnn
gcmc - cmc - m&c -

The materials used for a solar mission's heat shield or other hot structures should
combine a high temperature resistance and light weight, figure 6 22 illustrates the
relanonship between specific strength and temperature for various materials A more
i ed description of heat shield requirements and materials can be found in chapter

and eieva,ed — ~

carbon/carbon (C/C), mostly the chosen
temperature applications

material for very high

carbon/silicon (C/Si) for thermal protection

carbon/silicon-carbon (C/SiC), most suitable lightweight
structures at >1200 °C 6

material for hot

• tungsten

• refractory metals (low mass loss, high mass, brittle shells)

• ceramics (low mass loss, brittle shells, UV degradation)

• refractory composites (low mass loss, low mass, strong shells)

New materials are being developed [Bensimhon, 1996] [Randolph, 1996], e.g.:

titamum duminides to cover the temperature range from 650 °C to 850 °C
(beyond htamum capability) MU L

metal matrix composites: silicon carbide fibres in a metal matrix
• carbon fibre felts for thermal insulation

Carbon-carbon composites are the state-of-the-art material for

structures like heat shields, and they have been chosen for the heat shield of the SAUNA
spacecraft and the Suicide Probe. Typical material characteristics are [Ngai, 1991]:

• light weight and low density

. high strength and stiffness, which increase when temperature increases in
the range from 20 °C to 2000 °C

• low thermal expansion

• high thermal conductivity, decreasing with increase in temperature

• high thermal shock resistance

• high fracture toughness

• pseudo-plastic behaviour

• good fatigue and creep resistance

. controllable and predictable ablation, erosion, and recession characteristics

• excellent wear rate, applicable when a high coefficient of friction is required

Different types of carbon fibres are available. When structural properties are important
high strength, high modulus fibres are selected. High modulus fibres provide a high
thermal conductivity and low thermal expansion. If low thermal conductivity ,s
necessary, low modulus fibres are preferred.

Ceramic Matrix Composites (CMCs)

CMCs can be divided into non-oxide ceramic systems, which are silicon carbides (SiC)
and silicon nitrides. The oxide ceramic matrix system in use is alimuna The tensile a
flexural strength of SiC/SiC show a maintenance of properties up to 1500 C. Ihe
coefficient of thermal expansion of SiC/SiC increases for temperatures higher than
100 °C, which is critical for joints. CMC-materials can have long lifetimes.

Long manufacturing times and expensive raw materials lead to very high prices for the
finished CMC components, which restricts their application in space. To overcome these
restrictions, DLR in Germany is currently pursuing a low-cost technology [Krenkel et al
19951 If one day CMCs can more commonly be used in space, there might be new
conclusions. Currently, we see no advantage over carbon-carbon composites for a solar

mission.

Metal Matrix Composites (MMCs)

Typical metal-matrix materials are aluminium, magnesium, copper and titanium. MMCs
are used in jet and car engines where they provide high power to weight ratios. They
have the properties necessary for elevated temperature applications. These properties are
low densfty ,high specific strength and stiffness, high thermal conductivity, good fatigue
response, control of thermal expansion, and high wear resistance. But degradation of the
properties of MMCs still starts at about 300 °C. Still, we recommend the investigation of
MMCs for the propulsion system of solar spacecraft.

6.7 Thermal Control Technology Challenges

In this section the technological challenges, for the thermal control of a spacecraft which
travels toward the Sun, are discussed. Examples of such spacecraft are a flyby probe or a
suicide spacecraft. 3 3 F

The objective of a thermal control system is to control the temperature of the instruments
withm the required range. The thermal control can be subdivided into two parts, the heat
shield to obtain a shadow for the spacecraft and the thermal control of the instruments
For a spacecraft near to the Sun (2 -20 R s ) the use of solar arrays is difficult due to the high
thermal radiation flux. Therefore the possibility of using the heat for generation of
electric power is discussed.

6.7.1 Thermal Environment

The environment of a spacecraft, when it travels to the Sun, can be subdivided into:

• Radiation of the visible and the short-wavelength (50 - 140 nm)

electromagnetic radiation ' '

• Solar wind, the flux of particles ejected by the Sun

• Outgassing due to evaporation and desorption of the spacecraft materials, a
plasma is created around the probe

The radiative heat flux from the Sun (Qsun) received at a distance equal to the Earth's
mean orbital distance is known to be 1353 W/m 2 . Assuming that the Sun is a perfect
sphere and its radiance varies with spherical symmetry, the normal heat flux at distance
R (measured from the centre of the Sun) becomes,

Qsun = 6242.5 *10 4 «

(W/m 2 )

( 6 . 2 )

where R s is the Sun's radius being 6.96M0 s km [Park et al., 1981], The solar
electromagnetic radiation has a very wide spectral range. The long-wavelength
component contributes only a very small amount of heat and is therefore negligible.
However, they can significantly degrade the optical properties of the heat shield. In the
short-wavelength range (50 - 140 nm), the most prominent is the radiation at 121 6 nm
caused by the lyman-alpha line of the hydrogen atom. This radiation can ionise surface
atoms when it reaches a hot surface or gas and therefore can cause interference with the
scientific measurements onboard the spacecraft.

The solar wind consists of particles ejected from the Sun. At a distance less than 10 Rc
the behaviour of the solar wind is unknown. Hence, determination of the solar wind in
this near-Sun region is one of the scientific objectives. The particles can interact with the
heat shield material, changing the optical properties of the heat shield surface. The
optical properties determine the temperature of the heat shield. Therefore, the studying

of the effects of the solar wind on the optical properties of the heat shield material is
important [Park et al, 1981].

When a spacecraft approaches the Sun, the temperature of the heat shield will increase
Therefore mass loss will occur due to:

Sublimation of heat shield material generating a plasma of heat shield
material around the solar probe

• Outgassing of air molecules adsorbed while the spacecraft was in the
Earth's atmosphere

The outgassing species can generate a self-induced plasma cloud around the spacecraft at
a time when the instruments were attempting to measure natural plasmas around the
Sun. Thus, the principal requirement on the heat shield design is to minimise the mass
loss [Randolph et al, 1996].

6.7.2 Description of a Heat Shield

When a spacecraft travels towards the Sun the heat flux will increase dramatically
[equation 6 3]. This requires the spacecraft to be protected within the shadow envelope of
a protective shield. The temperature of a shield (Tshield) between the spacecraft and the
Sun is about:

Tshield = }

l Qsun a sin(0)
a e(F+l)

where: Qsun
A

$

6

e

F

= heat flux from the Sun

= Stefan-Boltzmann constant (5.728• 10' 8 W/ m 2 • K)

= solar absorptivity of heat shield material (typically 0.91)
= emissivity of heat shield material (typically 0.82)

= angle of incidence (approximately 30 °)

= backside view factor to space (typically approximately
0.7)

Using the typical values for a carbon-carbon heat shield the temperature for a distance of
0.2 AU is 600 K [figure 6.23]. This temperature can be decreased by decrease of the i/O
ratio and/or decrease of the angle of incidence.

2500

2000

1500

\

\

\

\

\

\

\

-Tshield [K]

-Solar flux

[W/cm /v 2]

Figure 6.23

Temperature of a typical heat shield and the heat flux as a function of the
distance from the Sun (steady state).

So the mam properties of the heat shield necessary to decrease the temperature of the
heat shield and the heat flux to the spacecraft are:

Low solar absorptivity, low $/6 ratio, which are optical material surface
properties.

High angle of incidence, which is restricted by dimensions of the launcher
and structural constrains.

In addition, the mass loss rate (dm/dt) of the shield at elevated temperatures must not
interfere the measurements of the plasma, other requirements are;

• No change in optical properties during the spacecraft lifetime, as these
determine the temperature of the shield.

• Minimal mass of the heat shield construction.

• Dimensions compatible with the launcher.

• No change in mechanical properties during spacecraft lifetime.

• The heat shield must be resistant against vibration during launch.

Previous studies evaluated the candidate materials, refractory metals (tungsten, rhenium,
tantalum, and their alloys), ceramics and refractory graphitics (graphite, carbon-carbon)!

™ S ( condl [ ded that carbon-carbon is the most appropriate material [Randolph et ai,
1996J for the heat shield due to:

• Low density and high elasticity modulis, therefore a low heat shield mass.

Technolncrv (TOC Ti

i crc

• Low vapour pressure, which results in a low mass loss rate.

• Stability of properties in charged particle and high ultraviolet flux
environments.

• Experience with the manufacture and characterisation of this class of
materials.

• Relatively stable ratio of solar absorptivity to emissivity (t/O).

The major reason to reject the refractory metals is the high density, brittleness and the so
called "darkening" due to charged-particle flux. This darkening decreases the ratio
($/6) and therefore increases the temperature of the heat shield. The major concern of
ceramics is the degradation of the material due to UV-radiation and not nearly as mature
in technology development as carbon-carbon.

The absorptivity of carbon-carbon is about 0.9 which is high, and at first consideration, it
would seem that a reflective surface (absorptivity small) would be an advantageous
material selection to minimise the absorbed radiant solar energy. Up to now this
approach is rejected because of the unknowns associated with the response of reflective
surfaces to the charged-particle, ultraviolet radiation and micrometeoroid fluxes that the
spacecraft is exposed to during the long flight time. Determination of the material
response to these environmental conditions is very difficult to simulate with ground tests.

In a previous solar probe study the requirement for the mass loss rate is less than

2.5 mg/s [Millard, 1992]. The requirement of 2.5 mg/s is bases on a flyby of the Sun
which takes only a few hours. However for a spacecraft which is in orbit around the Sun
the value of 2.5 mg/s results in a total mass loss of 10 kg (of a 10 kg heat shield) within

1.5 months. Therefore the requirement of 2.5 mg/s is not sufficient, decrease by a factor
100 up to 10 4 may be needed.

The mass loss rate may be predicted by the Langmuir-Knudsen equation, which shows
that dm/dt, for a certain material, can be decreased, by decreasing the temperature of the
heat shield. For carbon-carbon the dm/dt increases a order of magnitude for every 100 K
[Randolph et al., 1996] increase of temperature. Possibilities to decrease the temperature

of the heat shield are:

• A coating with industrial diamond powder.

• Surface treatments, chemical vapour deposition of pyrolytic graphite
coating [Randolph et al., 1996].

• Thin sheet of tungsten, which is protected against darkening during the
flight to the Sun by a thin protective layer. If the protective layer will
become too hot it will evaporate and the low absorption tungsten sheet will
decrease the temperature of the heat shield.

• Making the angle of incidence smaller, by using a larger heat shield. To
accommodate the shield in the launcher, a deployable shield can be used.

6.8 Guidance, Navigation and Control

In this section all the issues related to the functions to be performed by the spacecraft
system as a whole to know its position, velocity and attitude are included. The Guidance,
Navigation and Control (GNC) subsystem is responsible during the whole spacecraft
lifetime to maintain the required position and orientation during every phase of the
mission. The analysis and selection of the trajectories to be followed by the spacecraft is
performed in the section 6.3.

icc « i?*. TTio Qnn for Scipnep and Humanity

Both concepts, position and attitude determination and control, will be dealt
this chapter in the sections 6.8.1 and 6.8.2, respectively.

separately in

With a focus on describing challenges and possibilities for GNC advances in relation to

solar missions, this section will take into account possible enhancements in the following
fields of interest: 8

Increase the knowledge of the environment faced by spacecraft in an
interplanetary mission, in particular near the Sun and near other planets (if
fly or swing-by operations are required). Better modelling of this
environment will improve the operation of the spacecraft (from the point of

view of the GNC and ACS subsystems) because of better design of the
control system.

• Increase the performances of the existing sensors or development of new
measurement techniques.

• Increase the computer performances to allow more complex spacecraft
operations. In this area, the growth has been exponential during the last
years leading to a remarkable increase of the functions that the satellite can
perform in an autonomous way.

• Increase the performances of the existing actuators or development of new
actuator systems.

The current state of the art in this technological area will be briefly reviewed to use the
most advanced techniques available in the near future to solve the technological
problems derived from these missions with the Sun as objective. °

6.8.1 Orbit Determination and Control

Orbit determination and control commands are usually provided from the ground for
interplanetary missions and it is based on the tracking of the spacecraft by radio signals.
Triangulation techniques using on-board instrumentation are available but they have a
reduced accuracy in the position determination when compared with the ground-based
technique. These triangulation techniques may be based on measuring star or planets
directions or the time of occultation of some stars behind close planets or satellites (it
could be useful during flybys) [Battin, 1989]. The actions to be taken by the spacecraft to
correct the position are also commanded from ground, by using a propulsion system
(based on hydrazine, cold gas, etc.). y

Ground control for orbit determination has been extensively used for interplanetarv
missions and it has also been used for near Earth satellites until Global Positioning
System (GPS) became operational on 1989. Nowadays, on-board orbit determination
systems based on GPS signals are being analysed and planned for the next future.

™ bui 5 ftSf GPS system has had a strong impact in the current way to navigate in
LEO and MEO and it provides very exact position estimates. Perhaps, an equivalent
system to GPS but extrapolated to interplanetary navigation would represent the future
trame in this area if a system providing suitable performances is developed.

The development of such a system is out of the scope of this project with the Sun
exploration as the major objective, but some ideas about this concept can be highlighted

• The immediate and obvious extrapolation of the GPS concept to a Solar System
pomt of view would lead to a constellation of satellites orbiting around the Sun

providing a signal leading to a navigation solution in the area covered by these
satellites. This concept has strong problems due to the high number of satellites
required to provide a suitable coverage feature in a so huge area and to locate all
those spacecraft in orbits in different planes around the Sun to provide a suitable 3-
D measurement. GPS system is designed to provide four satellites in the field of
view of an antenna on the Earth surface during the 95 % of the time. The location
of the spacecraft in different planes around the Sun could require the use of new
propulsion technologies like ion propulsion, solar sailing, etc.

• Beacon spacecraft located in the L4 and L5 Lagrangian points of each Sun/planet
system. It would be a feasible solution for the mid-term future but it would
require a detailed analysis of the obtainable performances. A problem will be that
all the beacons will then be located in a near-ecliptic plane and only a 2-D
navigation solution will be available, being necessary an additional source of
information from out the ecliptic plane. Anyway, it could be a system to aid future
interplanetary navigation but it will not be autonomous by itself.

• Beacon spacecraft around each planet (or asteroids) of the Solar System or beacon
stations on planetary ground could be located to aid the navigation in the vicinity
of each planet.

Anyway, the development of such a system will only be interesting if an extensive use of
interplanetary flights will be done in the future when any ground control would be
overpassed by the high number of spacecraft. It is not justified at the current state of
spaceflight but in the future, the system could look similar to the current way used by
aeroplanes to fly, where the approach operations (flights around a planet) are controlled
by ground but the flight along the airways (interplanetary flight) is autonomously
performed by the plane itself. In this direction, the paper from Reidel may be interesting
to highlight the future autonomous navigation systems based on optical systems [Reidel,

1996].

As a summary, the orbit determination and control functions will be provided by ground
stations in the near-future. For the mid- and far-future, the system would be autonomous
using optical measurements navigation (mid-term) or by an interplanetary navigation
system (far-term).

6.8.2 Attitude Determination and Control

The ACS is the responsible to maintain the required orientation of the spacecraft due to
the need of pointing the solar arrays, the antennas, thermal control elements, the
instruments, the GNC propulsion system,...

A wide set of sensors and actuators are currently used providing a high number of
possible operational configurations in a near Earth operation.

The available sensors used by near-Earth missions are typically: Sun sensors. Earth
sensors, star sensors, inertial measurement units, GPS receivers and magnetometers. The
available actuators used in near-Earth missions are: magneto-torquers,

reaction/momentum wheels, thrusters and solar sailing.

Most of the commercially available sensors and actuators (for near-Earth orbits) have not
application to interplanetary missions. Therefore, only star sensors, gyrometers and,
maybe, Sun sensors could be used for interplanetary missions. Actuators which do not
require propellant consumption will be favoured because they will not limit the lifetime
of the spacecraft.

6.8.2.1 Environmental Issues

The definition of the environment to be faced by the spacecraft is a major task in order to
define the equipment selected for the on-board operations. A global view of the problem
provides the following set of environments:

Launch (common to every space mission and introducing requirements to
the structural characteristics of the sensors and actuators)

• Near-Earth environment (at least in the first stage of the mission)

• Interplanetary environment

• Environment of other planets (if fly-by operations are required and it could
include the planet atmosphere when aerobraking is used). Jupiter can
provide a very aggressive environment from the point of view of the
radiation due to its strong magnetic field (it is an issue to be taken into
account for all on-board electronic equipment).

• Near-Sun environment

Most of the environmental aspects are identified in the section 6.1. The attention will be
focused here in the environmental issues which lead to disturbances in the rotational
dynamic behaviour of the spacecraft. The usual elements which can introduce
disturbance torques to the system are typically planet atmospheres, gravity gradients,
magnetic interactions and solar radiation pressure.

Obviously, atmospheric disturbances are only taken into account when a fly-by at a low
altitude over a planet with an atmosphere is performed and during Earth orbit
operations. However, these operations are only developed during a short period of time
with respect to the total mission duration to take it into account as a major issue in the
ADCS design.

Gravity gradient disturbance is a torque produced by the no coincidence of the centre of
mass of the spacecraft and the centre of application of the gravitational forces acting on
each particle of it. These effects are not relevant during interplanetary flight phases and
they are not taken into account during fly-by operations or operation around L4 or L5
points, but it may be relevant for spacecraft orbiting the Sun if the distance is relatively
small and if it is going to operate at that distance for a long time. The values for SAUNA
has been estimated and are negligible when compared with solar radiation pressure
disturbance.

Magnetic disturbance torques may appear due to the residual magnetic field of the
spacecraft itself. The value of this field is difficult to predict but can be measured after
the spacecraft integration. It would be desirable to reduce it to a minimum because the
magnetic fields that the spacecraft will find are not well-known and it could increase the
operational problems of the spacecraft. Anyway, if we are going to have magnetic
measurements, the value of this residual field should be small.

Solar radiation pressure disturbance will have a major effect on the spacecraft dynamics
being actually the dominant effect. This disturbance appears due to the non-coincidence
of both the centre of mass of the spacecraft and the centre of the solar pressure force over
the spacecraft surface. This will produce an external torque on the spacecraft changing its
angular momentum. 6 6

When operation at L4 or L5 is considered, solar radiation pressure is still the dominant
perturbation but with a considerable reduced intensity because they are at the same

T

i r~r\

distance from the Sun as the Earth but a spacecraft operating on them will not have
disturbances from Earth atmosphere, gravity or magnetic fields.

6*8.2.2 Other Missions ADCS Review

To start the definition of this subsystem it is necessary as a preceding step to review the
past missions, in order to identify the technological problems that have already been
solved. The review should not be intended as an exhaustive recompilation of every
mission but the compilation of most relevant configurations that could guide our design.

For missions with the objective to operate as close to the Sun as possible, the paradigm to
take into account at the current state of the art is the Solar Probe mission, planned to
operate at a distance of 4 Rg over the Sun with a three-axis stabilised configuration.

The Solar Probe attitude determination system [Randolph, 1995] is based on star trackers
and an HRG (Hemispherical Resonator Gyro), which will be used alone during the
perihelion pass in order to avoid possible malfunction of the star tracker. The HRG drift
is periodically calibrated by using the measurement from the star trackers. The Cassini
HRG is still considered to be the only existing design capable of meeting the drift
requirements of the mission (the drift is 0.006 °/h, 3-sigma) with an internally redundant
capability and radiation hardness needed for the mission. The largest impact was the
realisation that the steady state power usage of the Cassini HRG is now estimated at
24.6 W, and it will be worse during calibration periods with the star trackers operating at

12 W. '

HRG technology advances that would allow for comparable accuracies at lower power
usage and less mass are under continuous assessment.. At this time there are no
available qualified designs with superior performance and lower power and mass
characteristics. It is anticipated that within two to three years an IFOG technology
(Interferometric Fibre-Optic Gyro) may challenge the HRG performance and have lower
mass and power levels. IFOG is planned to fly on Clementine [Eliason, 1994].

The Hughes Danbury HD-1003 unit is the star sensor used in the Solar Probe mission.
The accuracy of this instrument is about 50 prad.

The actuator system has evolved from the early versions of the Solar Probe to the last one
called Minimum Solar Mission (MSM) to become a single actuator system based on
hydrazine to perform both the orbital and the attitude corrections. The former version
based on a dual engine configuration with a cold gas system to perform the attitude
corrections has not been selected at last.

PLAMYA mission attitude determination [Randolph, 1995] is based on an IRU calibrated
by optical sensors for Sun and star measurements. The IRU contains four angular
velocity sensors, each capable of measuring rates up to 10 °/s. The angular velocity
measurement error will be less than 0.05% with a drift rate of less than 4.5 arcmin/mm.
The lifetime is expected to be 50000 hours.

The optical instruments are provided by the firm NPO "Geofizika [Randolph, 1995] and
the characteristics of them are summarised below:

• Sun sensors: 40 ° x 186 ° Field Of View (FOV) with measurement error not
exceeding 2 arcmin in the FOV. Redundant units are provided with total
mass of 6.7 kg and 8 W power consumption. These allow for drift
measurement to better than 0.003 °/hr and determination of the axes

■i /n

d~. r ru^ frt*. Crionro anrl Hnmanitv

orthogonal to the Sun to within 2 arcmin and roll error about the Sun line to
no more than 1.5 arcmin.

• Star sensors: bright star sensors to measure references from Sirius,
Canopus and Vega exhibit a measurement error of no more than 3 arcmin
within a FOV of 2.6 0 x 2.6 °. The instrument has a mechanical device
which can move the FOV around one axis with an error not greater than
2 arcmin, within a range of ± 37 °. With redundant instruments, the mass is
29.5 kg and power consumption is 21 W.

Another potential set of star sensors is the "stellar occultation instruments" (POZ)
[Randolph, 1995], which utilise certain parts of the stellar sky along the flight path The
firm NPO "ELAS" has an instrument with an 8 ° x 8 ° FOV. This device defivers angular
co-ordinates of up to 8 point objects located in the FOV with a brightness range of 0 to +6
stellar magnitude. The limiting total error of the instrument is less than one arcmin. A
solar protection cover is automatically closed when the Sun is near the FOV of the
instrument. Two sets of instruments have a mass of 11 kg, while the electric power
consumption of one instrument is 22 W. This device has not yet been flight tested.

The actuator system is based on thrusters and it is commanding by a reaction devices
control system. An amplifier/converter have a mass of 3 kg and consume 10 W of power.

The Magellan mission to Venus [Young, 1990] attitude determination system is based on
an IRU calibrated periodically by a star scanner, entering into a calibration mode with a
required inertial pointing. The gyroscopes are required due to the high number of
manoeuvres to be performed by the spacecraft. The control of all these operations is
performed by reaction wheels, reducing the amount of fuel needed for the manoeuvres
As a complement to the attitude control hardware, there is a set of Sun sensors and solar-
array drive motors which keep the solar panels pointed toward the Sun.

6.8.2.3 Review of Available Instrumentation

As it has been seen in the previous section, most of the designs are based on star trackers
and IMUs for attitude determination and reaction wheels and thrusters (cold gas and
hydrazine) for attitude control. They are the baseline instrumentation for this mission
The following table lists accuracy, FOV, and sensitivity data for some star sensors.

Table 6.19 Accuracy, FOV, and Sensitivity (mv) of Some Star Sensors.

Accuracy

FOV

mv

Star Tracker Stellar Compass (STSC) -
LNLL (LLNL, 1996]

150 prad (p&y)
450 ixrad (r)

28.9°x 43.4°

<4.5

OCA'S WFOV LLNL [LLNL, 19961

—

28°x 44°

4.5

CT-601/602 Ball [Ball, 1996]

3 arcsec

8°x 8°

+ 1 / + 6

CT - 611 Ball [Ball, 19961

15 arcsec

10°x 10°

-7/ +3.9

CT-621 Ball [Ball, 1996]

11 arcsec

20° x 20°

+ 0.1 / +4.5

CT - 631 / 632 / 633 Ball [Ball, 19961

20 arcsec

20°x 20°

+ 0.1 / +4.5

Mini Star Tracker

Clark Technologies [Clark, 19961

6 arcsec (p&y)
20 arcsec (r)

“*

NPO "Geofizika" Star Sensor
(PLAMYA) [Randolph, 1995]

< 3 arcsec

2.6°x 2.6°

Bright stars

NPO "ELAS" Stellar Occultation ]

instrument [Randolph, 19951

< 1 arcmin

8°x 8°

0/+6

Astro 1M [Elstner et ai, 19911

1 - 2 arcsec

5.3°x 8°

<8

The performances of star sensors will increase when the new concepts based on APS
(Active Pixel System) will be developed. Some information about this system based on
the use of CMOS technology can be found in [JPL, 1996].

The technology of gyrometers has evolved from the past gyros with moving mechanical
parts to the current designs based on HRG. The performances of HRGs are the best ones
at present time with a drift of 0.006 °/h. However, its main problem is its high power
consumption, but it is expected to reduce in the next future. New systems based on IFOG
are planned for the future with enhanced performances (Clementine, [Eliason, 1994]).

For missions located on L4 or L5, commercially available Sun sensors could be used,
because the Sun is seen in the same way than from Earth. For spacecraft orbiting the Sun,
it could be interesting to have Sun sensors. They should be based on a different concept
than the currently available ones to withstand the high temperatures and the high
relative size of the Sun. Maybe, this concept would not be possible for SAUNA mission
because it is too close the Sun.

In tables 6.20 and 6.21, the performances of some reaction wheels which can be used for
these missions are summarised.

Table 6.20 Performances of Some Reaction Wheels.

Performance

Speed

Ithaco

Ithaco

Characteristics

(r/min)

Type A *

Type B 4

Angular Momentum

1000

0.75

3.25

[N*m*s]

3000

2.25

9.75

5000

3.75

16.25

6000

N/A

19.50

Torque fmN«m]

20.0

40.00

Steady-State Power (W)

1000

3.0

3.0

3000

5.0

5.0

5000

7.0

7.0

6000

N/A

8.0

* [Ithaco, 1996]

Table 6.21 Performances of Some Reaction Wheels (Cont.).

Performance

Momentum Wheel
Ithaco *

Reaction Wheel
Ithaco *

Clark Technologies
**

Momentum Storage

80 N«m»s @ 6200
r/min

50N»m»s@3850

r/min

—

Max. Reaction

Torque

> 0.15 N*m

> 0.3 N*m

0.025

8 years

8 years

—

117 kg

14.1 kg

—

Power: Steady State
Peak

30 W@ 5000 r/min
110 W @ 5000 r/min

35 W @3850 r/min
200 W@ 3850 r/min

* [Ithaco, 1996] " [Clark, 1996]

No detailed information is provided about thrusters. They typically can be based on
hydrazine, cold gas or ion propulsion and the performances can be suited for the concrete
mission with the corresponding constraints.

162 • Ra: The Sun for Science and Humanity

6.8.2.4 Proposed ADCS Configurations

One important task leading the overall configuration of the spacecraft is the selection of
the type of stabilisation. Most of the reviewed missions are based on a three-axis
stabilisation in order to satisfy different requirements from thermal control
communications, instruments pointing, etc. Some spinning configurations have been
explored for the Solar Probe mission but none of them was selected. Three-axis
stabilisation has some benefits from the point of view of some subsystems when
compared with spinning options and the obtainable accuracy is potentially smaller
Dual-spin allows better performances than pure spin spacecraft but the problem then
appears in the lubricant to be used at high temperatures. Star sensors can only operate in
a suitable way when the angular velocity is under a relatively small value (~0.5 °/s).
Therefore, the spacecraft ADCS configuration would be based on a three axis
configuration unless spin options could be used depending on the mission.

Different equipment for attitude determination can be used. Star sensors provide an
attitude measurement independent of the position in the Solar System. If the mission
requires to perform some manoeuvres or if it includes a pass near the Sun, then an IRU
would be needed because the star sensor could not operate in a correct way in those
conditions. However, star sensors are needed to calibrate the drift of the IRU during
slow rate phases of the mission not close the Sun. For missions operating at L4-L5, one
star sensor head could be substituted by a Sun (it is a star too) sensor in order to reduce
cost.. Suicide probes would require an IRU to operate in their flight towards the Sun and
star sensors for calibration.

The actuator system could be based on reaction wheels (not limiting the spacecraft
hfetime) which need to be periodically desaturated using typically thrusters (solar sailing
could be an option). It could be advantageous to use the same propellant for attitude
control that the used for orbit control (this conclusion is true for Solar Probe mission
[Randolph, 1995]). Therefore, the selection of the attitude control could be guided by the
propulsion subsystem if it can be used for attitude control. For one-shot missions
(suicide probes), the control system could be based on thrusters alone because the
disturbance level is higher and the lifetime is not a constraint. This solution was selected
for Solar Probe because the estimated propellant mass to perform attitude control was
less than that required for the reaction wheel system.

6.9 Communications

This section is concerned with the design of the communications system for spacecraft in
interplanetary cruise phase and near-Sun environment. After a description of the
different issues related to communications, we have reviewed proposed technological
communication challenges to overcome these issues. The last part describes the
recommended approach for the implementation of the communications links for the Ra
missions in the near, mid- and long-term.

A lot of proposed or actual missions have either addressed or evaluated classical
problems, which have not been discussed here.

6.9.1 Communications Issues

The constraints on the data communication links in the context of the Ra project are
discussed hereafter. The discussion will be divided into problems related to special
conditions prevailing in the Sun's environment and those arising from mission

rpniimamonfc ° w

Technoloev Challpncyp*; Tccnoc •

6.9.1.1 Environmental Issues

Thermal noise

Thermal noise is a major concern in the vicinity of the Sun as the star appears as a noise
source producing interference at all frequency bands. This results in an increase of the
system noise temperature that has to be carefully taken into account when designing the
communications link.

The Sun's noise temperature depends on the frequency and the level of solar activity.
Several models are available for different bands [Maral et al, 1993] and yield values
varying from 200,000 K to 300,000 K at 1 GHz to 20,000 K at 10 GHz. The increase of
system noise temperature due to the Sun s contribution has to be weighted by the
receiving antenna pattern. The design should ensure that there is enough separation
between the Sun's direction and the antenna's main lobe, first and secondary side-lobes.

According to recent JPL measurements from a DSN station with a spacecraft close to the
Sun, at about 1.12 Rg (where the angle between the spacecraft and the Sun as viewed from
the receiving Earth station, or SEP angle is about 0.3 °), the system noise temperature in
X-band (8.4 GHz) ranged from 1417 K to 2300 K, depending upon the direction. In Ka-
band (32 Ghz) it was in the range of 456 K to 614 K. For antennas with a reasonably high
gain, e.g. 30 m at X-band (65 dB gain), thermal noise increase due to the Sun is no longer

a major problem.

Scintillation

As the spacecraft moves through or behind the solar corona, the relative geometry
between the transmitter, propagation media and receiver changes, and the received
signal will fluctuate like the twinkling of the stars. Similar to the ionospheric scintillation
that has been discussed in section 4.5, the plasma irregularities, or "blobs", will randomly
modulate both the phase and amplitude of the signal leading to significant degradation
of the link. This effect which is critical mostly around the solar corona is a major concern
for communications, but scientific information can be extracted.

Theoretically, the frequency dependency is /"' 5 at 1 GHz and above. For example, if
scintillation at 2.2 GHz is 6 dB peak-to-peak, then the scintillation at 7.25 GHz is 1.0 dB
peak-to-peak. Therefore the link degradation due to scintillation can be limited through
the use of higher frequencies.

A further understanding of the effects of scintillation will be gained through interesting
measurement opportunities offered by the NEAR mission. As the spacecraft will cross
the ecliptic plane, it will encounter a blob region and suffer from signal strength
degradation due to scintillation that can be monitored on Earth [Randolph, 1996].

6.9.1.2 Mission Related Issues
Spacecraft Configuration

The selection of a spacecraft configuration and of the communication scheme affect each
other mutually. Thus, in the case of a 3-axis stabilised spacecraft orbiting around the Sun,
the antenna needs to be continuously steered to be kept pointed towards the Earth. In
case the antenna is mechanically steered the reliability is significantly degraded due to
the use of moving parts. Therefore, alternatives need to be considered for cases where
the lifetime needs to be long or where no risks can be taken.

-I /-A

n „ . Tl,^ Qnrl T-Ti im tk nifv

Orbital Considerations

The problem of solar conjunction can be avoided by careful choices of trajectory and/or

c attx?a N USe ,i°L on ‘ board stora ge of data while in conjunction. Some Ra missions {e.g.
SAUNA) will have repeated solar conjunctions with a requirement for continuous high
data rate transmission throughout that time. When the spacecraft is behind the solar
disk, transmission is deemed impossible unless relay satellites are used or the spacecraft
have inter-satellite links. These will be discussed below.

For missions closer than 4 R s (e.g. Suicide Probe), there is a 1.5 0 arc field of view from
Earth for the Sun within which it is extremely difficult to use RF in a conventional
manner. Some missions will get so close that the phenomena observed will have a direct
e feet on the communication medium, e.g. coronal phenomena will prevent any kind of
radio emission. Other techniques will have to be investigated.

Interplanetary Travel

The link will have to be engineered to cope with the restrictions imposed by
interplanetary travel, e.g. trade-off between data rate and long distance/low power
related to Jupiter gravity assist.

Ground Segment

An early warning mission would require a constant coverage while a scientific mission
could do with on-board storage. If a network is considered, connectivity and accessibility
issues must be addressed.

6.9.2 Technological Communications Challenges

The different technological challenges that can overcome the issues presented above are
described here.

6.9.2.1 Radio Frequencies

RF communications have been used extensively - almost exclusively - in deep space
applications. It is a known and mature technology and as such, is a prime candidate for
low-risk missions, assuming the environment allows for transmission or the
distance/interference does not result in impossible power, antenna size and /or overall
mass requirement.

Figure 6.24 presents the advantages/disadvantages of the Ka band over the X band that
are the two frequency bands that we have considered for Ra.

Technolo^v Challpno-pc Rr Tcciinc a 1

Advantages
high data rate

impervious to scintillation

low power requirements

reduced hardware size

Disadvantages
atmospheric sensitive

not technically proven

poor ground segment availability

Figure 6.24 Advantages and Disadvantages of Ka Band Over X Band.

Due to scintillation the X band is not a favoured candidate in the vicinity of the Sun (< 4
R s ) except for particular geometrical configurations ( e.g . the Solar Probe [Randolph,
1996]). Therefore, we propose to limit the use of this band for up and downlink
communications while the Ka band is more suited for inter-satellite links.

6.9.2.2 On-Board Hardware

The typical block diagram of a deep space X band transponder is given in figure 6.25.

High

BLOCK DIAGRAM OF PROPOSED SPACECRAFT TELECOMMUNICATION SVSTEM

Figure 6.25 Typical Block Diagram of a Deep Space X Band Transponder.

6.9.2.3 RF Ground Segment

The design of the ground segment is a trade-off between the ground station and
spacecraft complexity. A typical block diagram of a ground station is available in the
appendix C.7. So far, most deep space missions have required full time coverage by high
gain antennas. Thus the Deep Space Network was a prime candidate, if not the only one,
to support such missions, leading to problems of availability of the network facilities.
Instead, the use of smaller stations should be considered and possibly make use of the

i aa • Ra* Thp Sun for Science and Humanity

availability of several US Department of Defence decommissioned antennas and/or
Russian facilities. Other possibilities include the use of antennas belonging to small space
agencies or organisations (e.g. GSOC of DLR), or the ESATRACK network of ESA. The
Villafranca station, currently used for the Infrared Space Observatory (ISO) mission has
the right characteristics for Ra [Maldari et al.,1996].

Continuous coverage should not be a driving factor in the ground station design but
instead should be the result of a trade-off between on-board storage, data volume link
capacity, and ground station costs.

6.9.2.4 Optical Links

The use of optical frequencies allows for a dramatic increase (> 90 dB) in power
concentration of the beam onto the target receiver compared to RF systems. Thus, optical
links can trade off some of that gain towards smaller aperture terminals, reductions in
power, size and mass, and increased link capacity. For a properly selected wavelength,
and with accurate filtering and pointing, reduction in background noise or direct
illumination from the Sun can be countered.

Direct detection might simply not be efficient enough to overcome the background noise
even though the transmit frequency is carefully selected. The amount of signal energy
that would have to be transmitted from the spacecraft would be unmanageable at such
distances from the Sun. Coherent detection seems more promising since light from the
Sun is inherently non-coherent. Thus the effect of background noise is acceptably

reduced. Heterodyning reception has the disadvantage of being more complex to
implement. r

In order to achieve the total power required, an array of diodes is necessary. Some diode-
pumped lasers are currently in use which produce power in the order of a few watts, vet
it is still a state-of-the-art technology. However, the power and reliability of laser sources
has been doubling every year and this trend does not seem to be stopping. Consequently
it is reasonable to assume that the necessary technology in terms of power rating and
lifetime will be there for some of our far-term missions (e.g. Suicide Probe) and we
recommend that this should be a focus for technology development.

Since spatial phase coherence has to be preserved, detectors should be ideally put in
space or the effects of the atmosphere would have to be taken into account. It is
envisioned that spaceborne reception will eventually be used [Lesh, 1992]. First it gets
the receiver above the cloud cover, as well as the phase front disturbances cause! by
atmospheric turbulence. Second, by being outside this same atmosphere, background
light associated with daytime scattering will be eliminated.

Spaceborne reception could become a reality in low Earth orbit, perhaps aboard or near
the Space station, by the turn of this century. However this still leaves Earth blockage to
a deep space probe, approximately half of the 90 minutes station orbit period if not more
due to Sun blockage when orbiting around it. In addition, it makes telescope pointing
more complex. Thus one would like to have such a station located on a much higher
orbit, perhaps geosynchronous, or at one of the stable libration points [Lesh, 1992], Once
received by the orbital terminal, the data would subsequently be relayed to the ground
via conventional RF techniques. The use of the Hubble Space Telescope has even been
proposed [Ashford, 1996]. A space-based interferometer could be constructed on either
side of the Earth, in Lagrangian points Earth-Moon, or using the south pole of the Moon.

If the receiving detectors are placed on the Earth then atmospheric effects must be
considered. Absorption of light from the atmosphere in the UV, optical and IR spectra

varies and is a very serious consideration since the total amount of energy received is
limited. Atmospheric turbulence is another effect to include in the optical link budget.

6.9.2.5 Inter-Satellite Links

Depending on the configuration and requirements of the mission/constellation (SAUNA,
Early Warning) inter-satellite links (ISL) will be required. As discussed earlier, the Ka
band is a suitable candidate for ISL if RF technology is considered. However, in the
vicinity of the Sun the use of laser for ISL is recommended. Solar interference, in terms of
beam disturbance and/or background noise, can be reduced with high pointing
accuracy. Depending on the data rate requirement ( e.g. > 1 Mb/s), optical links may be
the only cost-effective and technical solution available.

6.9.2.6 Advanced Antennas

Due to the extreme conditions of the near-Sun environment, a high gain mechanically
steered antenna is subject to thermal conditions that will limit the lifetime of the reflector,
joints, lubricants etc., on top of normal wear and tear. This reduces the system reliability
by introducing critical elements subject to single-point failure.

Other alternatives to the pointing mechanism are:

• reflectors with multiple feeds

• electronically steered phased arrays

The use of parabolic reflectors with multiple off-axis feeds is interesting to compensate
the motion in the orbital plane with respect to the ecliptic. These antennas have,
however, high mass and volume. In addition, antennas scanned off-axis have high losses,
even though for small displacements (in the case of SAUNA ±7 °) a shaped secondary
reflector can be used to compensate for these losses.

A better approach would be to use a phased-array antenna. Phased arrays steers
electronically the antenna by means of varying the phase or amplitude of each radiating
element. This would reduce the mechanical and structural requirements on the
spacecraft, allow for higher gain and possibly adaptive nulling of nearby interference
sources/ like the Sun.

For the SAUNA mission, the physical shape of the array would have to allow for a near¬
circular shaping of the beam, as the spacecraft will be orbiting the Sun and will have to
stay in permanent communication with Earth. Since all elements of the array are active,
the power required increases but, for SAUNA, this impact on the power budget would
not be significant.

6.9.3 Recommended Approach

This part presents the recommended technical approach concerning the communication
system for the variety of Ra missions and for the time frames considered.

6.9.3.1 Near-Term Programme

Technology development: Ka band transponder and ground stations, high power laser
sources, advanced antennas for deep space/Sun environment.

Missions: Concentrate on the existing technology, both in the space- and ground-
segment. Thus, the use of X band is recommended.

i^o A P-,. Tho Qnn fnr and HlimanitV

6.9.3.2 Mid-Term Programme

Technology development: phased array antennas to avoid the use of 2 degrees of
reedom mechanisms, space-qualified high power lasers and coherent detection
techniques, spaceborne detectors.

SAUNA: Use of Ka band in missions near to the Sun. This will yield an increase in
capacity, decrease in power, mass and size of the on-board hardware and antenna

SOLAR EARLY WARNING: The use of X band is foreseen for this mission as the
constellation is to be sited at 0.5 AU from the Sun. However the implementation of a
global network of ground stations is necessary, as the current capacity of the
conventional networks (e.g. DSN) is limited and would not be suitable for continuous
monitoring of the spacecraft.

6.9.3.3 Far-Term Programme

Technology development: Implementation of spaceborne optical receivers network,
leading to spaceborne interferometry.

SUICIDE: Communications for the proposed Suicide Probe will have to be implemented
using optical links. The mass and power constraints of the probe prevent any practical
use of RF. Moreover, the scintillation effects as the probe nears the Sun would simply
overwhelm the RF signal. It is hoped that an optical coherent-detected link, at a carefully

chosen wavelength, will be possible however, specific technologies will have to be further
developed.

6.10 Command and Data Handling

This section discusses various aspects of electronics, command and data handling for
solar missions. We focus on the following selected themes: °

• On-board electronics: Thermal and radiation environment.

Telemetry Processing: Standard telemetry formats, multiplexing

• Autonomy: Vehicle management, fault detection, isolation and recovery
(FDIR), payload data (pre)processing, collision avoidance, etc.

6.10.1 On-Board Electronics

The environment in the vicinity of the Sun is very harsh and extreme, especially in terms
of temperature and radiation. This means that the electronic equipment on-board the
spacecraft have to use a technology suitable to withstand and to survive this
environment. This is especially important when missions of relative long duration in
orbit and at a short distance to the Sun are considered.

In the Ra framework, missions of several years in orbit around the Sun at a distance of
around 30 Rg are proposed. This forces one to consider issues as reliability and protection
against degradation due to temperature and radiation. In addition, mission where
suicide probes are sent into the Sun corona are proposed. For this kind of missions, a

maximum survival time is desired, which poses additional requirements on the
technology.

Technology Challenges & Issues • 169

The high demands in terms of propulsion associated with getting close to the Sun
importantly constrain the mass available. This means that all the different elements on¬
board the spacecraft are highly constrained in mass.

In order to reduce mass associated to the electronic equipment, it is desirable first of all to
apply miniaturisation in order to reduce the mass and bulk of electronic components. In
addition, to reduce the mass required for shielding the equipment against temperature
and space radiation, it is very desirable to develop electronic technologies able to
withstand the high-temperature and high-radiation environment. Developments are
under way in space and other fields, which probably will provide sufficient demand for
the technologies to be mature enough for its use in space in the mid- or the long-term.

6.10.1.1 Miniaturisation

It is very desirable to reduce the mass and volume associated to the electronic equipment.
This leads directly to a significant reduction of cost or enables the integration of more
equipment within the same configuration. An example of the gain achievable with
miniaturisation is the unfortunate Cluster spacecraft, which development started in 1986.
In the 10 years gone by since then a substantial change in technology has been produced.
Thus, using technologies available today, the on-board data handling equipment would
have a mass at least one order of magnitude smaller. The application of micro- and
nanotechnologies to space applications for future space system goes far beyond and
predicts 40 cm microlanders with a mass of 550 g and 5 cm free-flying magnetometers
[Martinez de Aragon, 1995]

In the case of the electronic components this points out to the use of high-density
processes and advanced packaging technologies (e.g. high-density 3-D packaging) in
order to minimise the mass of electronic components. This advances in electronic
component miniaturisation will shift the critical point in electronic equipment
dimensioning towards the interfacing accessories (connectors and cabling) and
mechanical fixation to the spacecraft structure. Development in these areas will need to
be carried out in the far term in order to make use of the advances achieved and foreseen
in the near and mid-term.

6.10.1.2 Temperature Considerations

Even though there are possible missions in the Ra framework that would be exposed to
extremely low temperatures, the important concern is regarding the high temperature
environment near the Sun.

The approach followed in the missions carried out or proposed to date consists of using
conventional electronic technology in combination with important thermal shields to
keep the electronics and sensitive material at a reasonable temperature. This, however,
constrains the mission budgets, mass, volume, and hence cost, and imposes limits to the
spacecraft resources and configuration. Thus, for example, the NASA Solar Probe
provides a thermal shielding that assures that the electronic equipment don't exceed a
temperature of 40 °C during the perihelion pass (4 R*), where the maximal temperature is
reached [Randolph, 1996].

jfj thg near term and even in some cases in the mid-term, where budgetary or risk
constraints are imposed, mission planners will very likely have to use this approach for
missions in the vicinity of the Sun.

However, research is being carried out in different areas (automobile, aircraft,
communications and radar, power, and also spacecraft) on the development of new

17fl • Ra: The Sun for Science and Humanity

materials for advanced semiconductor electronic devices capable to withstand hostile
environments, and especially high temperature, high power, and high radiation

the NASA Lewis Hi s h Temperature Integrated Electronics and Sensors
(HTIES) Team is working is developing silicon carbide (SiC) as a material for these
demanding applications [HTIES Team, 1996], Even though the technology at this point is
immature, requiring improvements in crystal growth and device fabrication processes,
the enabling technology is available for it to evolve to meet the system demands for
hostile-environment electronics [Neudeck, 1996].

Silicon carbide electronics can operate at much higher temperatures (up to 600 °C) than
silicon (up to 125 °C) or gallium arsenide (limited to 350 °C). Therefore, the size and mass
o radiators and thermal shields on a spacecraft could be greatly reduced. This would
enable substantial mass savings on the spacecraft, or at least allow greater functionality
by utilising the volume and mass formerly occupied by the thermal management system.

urthermore, SiC electronic devices have also been shown to be less susceptible to
radiation damage than correspondingly rated silicon devices.

A more mature and today more cost-effective alternative to SiC is silicon on insulator
(SOI), even though its lower limit temperature makes the technology less attractive for its
use in extremely high-temperature environments. SOI devices withstand temperatures

up to 225 C for an operating lifetime of 5 years, and up to 300 °C with reduced operation
lifetime [Swenson, 1996]. *

e use of these high-temperature materials is very important for missions targeted to
the inner planets, where high temperatures will be encountered. So far, only very limited
use has been made of SiC for space applications. There are discussions in NASA about
the use SiC for Venus missions. However, the reduced number of missions planned in
the near future to the inner planets would not probably justify the investment required
for such a technology development. Nevertheless, the other applications, and especially
the automotive, aircraft and power industries will probably provide the pull required for
this new technology to develop and mature to an extent sufficient for its use in space
applications. Thus, even though it is unlikely that this technology will be available for
space applications in the near- or mid-term, it is expected that far-term missions aimed at
the inner planets or the Sun vicinity will possibly use SiC electronic technology.

6.10.1.3 Radiation Considerations

From the point of view of radiation, it is expected that the radiation fluxes in the Sun
vicinity will be very important. In addition, the long duration of some of the possible
missions proposed lead to very important cumulated radiation doses. Thus the
e ectronic equipment in mission scenarios with high exposure to radiation need to be
protected against the effects of this extreme environment.

Space Radiation Effects On Electronics

In space, high-energy particles can penetrate devices and cause temporary upsets and
permanent damage. Particle sources in space, and in the vicinity of the Sun in particular
are the cosmic-ray background, solar flare events and the solar wind. In planets with a
strong magnetic field, there are important number of trapped particles. Interactions with
spacecraft components cause secondary particle emissions as well. Finally, components
on the spacecraft itself, such as radioactive heaters, radioisotope thermoelectric
generators, and nuclear reactors, can emit particles.

In space, radiation effects on semiconductor devices are classified into two major types-
total ionising dose (TID) and single-event effects (SEE). TID is the accumulated effe^ of

ionising radiation over the lifetime of a space mission and depends not only upon total
trapped charge, but also upon the rate of incoming particles. SEEs are transient upsets
(soft errors, or single-event upsets, SEUs) or permanent damage (hard errors or latchups)
due to single particles. A third type, displacement damage, is less important [Messenger
et al, 1986] [ Rasmussen, 1988] [Stassinopoulos et al, 1988].

Soft errors (single-event upsets in storage elements and multiple-bit upsets in some types
of memory devices) are temporary since they merely cause a logical error, they do not
damage the chip. Once any affected registers are reset, the chip will resume correct
operation. But that error has the potential to propagate and effect critical functions,
causing any number of permanent problems on board the spacecraft. As an example, a
soft error in a register of the critical control electronic elements for deployment or attitude
control may initiate potential catastrophic failures.

Finally, the high number of particles and highly energetic environment could cause
interference and charging problems. In order to avoid such problems a well planned
grounding scheme together with interference mitigation measures must be engineered.

Table 6.22 summarises the components of the natural space radiation environment and its
primary effects in CMOS devices, by far the most used electronics technology.

Table 6.22 Summary of Space Radiation Environment and Their Effects on CMOS
Electronic Devices.

Radiation

Source

Particle

Types

Primary Effects

Solar wind

Trapped radiation belts

Electrons

Protons

Ionisation damage

Ionisation damage; SEE in
sensitive devices

Galactic cosmic rays

High-energy charged particles

SEE

Solar flares

Electrons

Ionisation damage

Protons

Ionisation damage; SEE in
sensitive devices

Lower energy heavy-charged
particles

SEE

Protection Strategies Within the Radiation Environment

Radiation shielding is an integral part of any spacecraft design. The best shields have low
atomic number, such as carbon and aluminium. Shielding can significantly reduce TID,
but it can rarely affect SEEs since particles energetic enough to cause SEEs typically
require shields several inches thick to be adequately attenuated. Unfortunately, shielding
may also enhance TID and SEEs by slowing fast particles into energy ranges of SEE or
TID sensitivity.

Given that flight path considerations and shielding cannot completely shelter electronics
from radiation, designers must use radiation hardened (rad-hard) or radiation tolerant
electronics, depending upon the radiation total dose and flux and fault tolerant
subsystems as the final recourse [Kerns et al., 1988]. Radiation tolerant electronic
components normally can withstand up to a few tens of krad(Si), while rad-hard
components withstand hundreds of krad(Si) or up to Mrad(Si). Rad-hard electronic
components are normally manufactured in CMOS-SOS technology, even though some

other hardening technologies exist that have been used occasionally, such as epitaxial or
sihcon-on-insulator (SOI) substrates or bulk CMOS.

Fault tolerance includes built-in self tests, redundancy, and other methods. Built-in self
tests constantly check faults so that the system can implement backup procedures
immediately. Redundancy can be implemented in different manners, depending upon
the requirements on reliability and the amount of risk acceptable,

• Hot redundancy, which consists of several elements operating in parallel,
with additional components devoted to deciding on correct results, such as
majority voting schemes.

• Cold redundancy, where the redundant equipment are switched off until a
supervisory circuitry detects a failure in the nominal equipment. Then the
function is taken over by the redundant equipment.

One frequently used approach to harden a system against SEU effects is to apply error
detection and correction (EDAC). ED AC can be implemented in a number of ways, and
can be a very effective way to accommodate SEU-induced errors in memory
microprocessor, or interface blocks.

The problems associated with the use of rad-hard electronic components are a
comparably lower density of integration and substantially higher power consumption.
This results in significantly higher mass and power budgets. This fact has an important
impact for missions where large amounts of electronic components are required. A clear
example is in solid state recorders (SSR), where large quantities of memory are required
(in the order of hundreds of Mbits or even Gbits). The bulk and power consumption
associated to the use of rad-hard technology in this case, very likely rules out the
possibility of rad-hard SSRs. Instead, the combined use of shielding, redundancy failure
detection and recovery (FDIR) and EDAC together with high-density memory devices
and advanced packaging is foreseen for such recorders, even in harsh environments
[Seidleck et al., 1996].

6.10.1.4 Recommended Approach

Given the extremely harsh environment in the vicinity of the Sun, and considering the
state-of-the-art and the foreseen technology development in electronics for high-
temperature and high-radiation environments, we suggest the following approach for on¬
board electronics:

Near-Term Programme

• Standard CMOS technology for non-critical electronics with adequate
thermal and radiation shielding, FDIR and EDAC

• CMOS-SOS technology for critical electronic components with adequate
thermal shielding, FDIR (redundancy)

Mid-Term Programme

• SOI or CMOS technology for non-critical electronics with adequate
(reduced) thermal and radiation shielding, FDIR and EDAC

• CMOS-SOS technology for critical electronic components with adequate
thermal shielding, FDIR (redundancy)

• SiC technology depending on technology maturity at the time

Tprhnnl r\cr\r T.

Far-Term Programme

• SiC technology for most electronics with adequate (highly reduced) thermal
and radiation shielding, FDIR and EDAC (if at all needed)

• Standard CMOS technology for non-critical electronics requiring high
density or high performance, with adequate thermal and radiation
shielding, FDIR and EDAC.

6.10.2 Telemetry Processing

The following section will discuss telemetry processing. Two topics that will be
discussed are standard telemetry formats and multiplexing techniques.

6.10.2.1 Standard Telemetry Formats

The use of a standardised telemetry format such as the CCSDS format can contribute to
reductions in ground efforts, e.g. in mission control centre software and in the world¬
wide utilisation and processing of the spacecraft data. Such standardised formats are
advisable for science missions [SAUNA, chapter 9.1] to allow total format compatibility
and easy access to the data.

A standard telemetry format is absolutely required for larger constellations of spacecraft
such as the solar environment monitoring networks discussed in the far-term time frame
of the Ra Strategic Framework; without it, each spacecraft would require its own software
at the operations centre, and this would make ground support very expensive for such
networks.

6.10.2.2 Multiplexing Techniques

Multiplexing is the process where multiple channels are combined for transmission over
a common transmission path. There are three predominant ways to multiplex (hybrids of
these techniques also exist):

Frequency Division Multiplexing (FDM)

In FDM, multiple channels are combined onto a single aggregate signal for transmission.
The channels are separated in the aggregate by their frequency. Signals occupying non¬
overlapping frequency bands are added and any one of these can be recovered by
filtering.

Time Division Multiplexing (TDM)

In Time Division Multiplexing, channels "share" the common aggregate based upon time!
Signals are compressed into high speed bursts which are placed in non-overlapping time
slots within a time frame. Recovery of the original burst is accomplished by selection of
the specific time slot in which the burst is positioned. Clearly this procedure requires
timing references.

Code Division Multiplexing (CDM)

With code division multiplexing all users simultaneously operate within the same
frequency band and each user occupies all the time the entire transponder bandwidth.
Each user combines the signal to be transmitted with a signature sequence which
displays two main correlation properties: (1) each sequence can easily be distinguished
from a time shifted version of itself; (2.) each sequence can be easily be distinguished

174 • Ra: The Sun for Science and Humanity

from every other one in the set. Using these properties the receiver is able to separate the
received signals even though they occupy the same bandwidth at the same time.

There are many solutions to the problem of multiplexing to a repeater by a group of
network stations. The choice of access type depends above all on economic
considerations: there are the global cost in terms of investment and operating cost and
the benefits in terms of revenues [Maral ef a/.]. ' °

Multiplexing may be pushed to the limit of current performance capabilities by the
advent of large satellite constellations (as proposed in the Strategic Framework)
providing early warning data through a lesser number of Earth relay satellites.

6.10.3 Spacecraft Autonomy

The Ra missions require significant advances in technology as the programmes outlined
in the Strategic Framework [chapter 2] become more and more ambitious with time. The
increased complexity makes the spacecraft difficult to operate and also very dependent
on the correct operation of all involved instrumentation. More on-board systems shall be
integrated on the same computer and utilise the same instrument. This integration is
demanding for the design process and the on-board autonomy. This section suggests
improvements on different operational aspects related to on-board autonomy, which is
feasible with increased computational power of future spacecraft.

6.10.3.1 Rationale for Autonomy

The two main objectives for on-board autonomy are to decrease the cost and to improve
the performance of the spacecraft without increasing the risk. Several aspects related to
these factors are displayed in figure 6.26 and described in the sequel. Some guidelines
are specific to the individual Ra missions, but all shall be applied in some degree. A large
variety of techniques can be used to increase a spacecraft's level of autonomy. More on¬
board automation requires more processing, but several advantages are possible.

Techniques

Replace hardware
by software

1 C

\ ^On^oar^uWgation

$ / -

v I Command validation

Science data
selective compression

£

C

C

Science operations
management

Intrainstrument

communication

c

Autonomous

retargetting

Multiple spacecrafts
intercommunication
and coordination

^^na^coorainatioi^

>acecrafte\ f #eUer

imication I V I of .

continuity

science

Figure 6.26 Factors Relating to Spacecraft Autonomy.

The ground operational expenses can be reduced by moving some of the functions
traditionally performed on the ground to the spacecraft and also by reducing the
requirement for a network of tracking stations. One example is to have on-board
orbit/trajectory calculation, so periodic upload of parameters is avoided. An approach
for interplanetary navigation has been proposed by [Bhaskaran, 1996] based on optical
navigation using asteroids as beacons. This is, of course, only applicable in regions with
asteroids that have known ephemeris data. The Ra missions operate inside the Earth's
orbit with almost no asteroids, so the principle is not applicable. Instead, we propose to
develop the concept to operate on planets. The principle is to measure the angle between
a near object (a planet) and an inertially fixed object (a star). Applying multiple
constellations, three axes position determination should be possible in the inner solar
system. The SAUNA mission [section 9.1] needs on-board orbit information to control
the pointable antenna. An on-board orbit model can provide this, but automatic
navigation is desirable because a model has drift errors and needs updates from ground.
If the antenna pointing is wrong, the spacecraft life depends solely on the low gain
antennas. Anyhow, it may not be possible to develop the technique within the SAUNA
programme, because the budget assumes mostly well known technologies.

The Ra mission satellites are designed to automatically reconfigure in case of anomalies,
because there will be periods without ground contact and when contact is possible, the
communication round-trip is up to the order of an hour (8 minutes for 1 AU). The
Suicide Probe [section 10.2] goes as far as 5 AU from Earth, so real-time decisions must be
taken on-board. Fault detection, isolation and recovery (FDIR) has the objective to
provide a graceful degradation in case of minor anomalies, so the maximum performance
of the spacecraft is utilised so as to keep the payload in operation as long as possible and
enter a safe mode if the spacecraft health is in danger. Modem methods using analytical
redundancy (exploiting the relationship between the input and the output of a dynamic
system) in combination with advanced statistical methods will be applied in the Ra
missions and thereby reduce the level of sensor redundancy.

On-board command validation is considered for all Ra missions, but it is specifically
important in the multiple spacecraft missions [Suicide Probe, section 10.2; Early Warning
System, section 9.2]. Commands and information transmitted from one spacecraft to
another can be erroneous, but wrong commands can also be sent from ground, because
international network operations have people employed from different nationalities and
ground personnel may be renewed. Ground validation shall also be performed to the
extent possible, but with increased spacecraft autonomy, the ground station may not
comprise all required information.

The communication system in all the Ra missions is critical because of limited
transmission power and mass. Therefore, science telemetry data rate will be reduced by
data compression and by science operation management control. Modem methods for
compression can be implemented in either hardware or software and make a significant
reduction with none or very little loss of information. On-board selection between
different compression methods (or bypass) can be installed to match specific data
sequence characteristics. Science operation management comprises the triggering of
special modes of the science instruments, like high speed data acquisition, activation of
measurements, attitude manoeuvres for targeting, and eventually ejection of the Suicide
Probe from the mother spacecraft. The ability for communication close to Sun is
uncertain in the present design (SAUNA will probably not be able to have high data rate
in 2/3 of the time), so it is very important to implement an intelligent manipulation of
science data in case of reduced downlink capabilities, including prioritisation of science
categories.

176 • Ra- TV»a Sun for Science and Humanity

The Early Warning System proposed in section 9.2 with 20 small spacecraft floating
around the Sun provides some interesting possibilities for increased autonomy between
he individual satellites. Intercommunication can provide synchronisation of actions or
even remote diagnostic of spacecraft behaviours with respect to anomaly detection. It is

also possible to implement the measurement analysis on-board, so only a detected alarm
is sent to ground.

6.10.3.2 Previous and Planned Missions with Focus on Autonomy

Autonomy is not a new concept. Some of the spacecraft flown to develop different

l° mat i? n are UlySSeS ' Clem entine, TAOS, and XTE. Future missions include
SA s PROBA satellite, NASA's New Millennium Programme, and the Japanese MUSES-
C satellite proposed by the Institute of Space and Astronomical Science. These do not
constitute a complete collection of spacecraft designed with the attribute of autonomy,
mainly because the concept of autonomy has different interpretations. In this context
autonomy is considered as decision making, a higher level of automation than signal
processing and feedback control loops. Further information can be found in the

iAnAn7; in ^i referenCeS ' Usted in ° rder of a PP earance; [Ulysses Spacecraft Home Page,
WWW; Clementine Information Home Page, WWW; TAOS, 1996; Technology for

Autonomous Operational Survivability (TAOS) Satellite Home Page WWW' Day et al
1996; Francesco, 1996; Lisman, 1996; and Nakatani, 1996].

6.10.3.3 Design and Implementation Issues

The implementation of more autonomy suffers from the paradox that increased
complexity also raises the inherent probability of failures. Therefore, generic methods
shall be developed and used to ensure completeness and correctness of on-board
decisions. Tools have been developed by the Artificial Intelligence community that can
assist to organise the inter-relationship between a large number of on-board functions. It
is very important that the end-product has improved reliability, so the operator does not
just disable the autonomous functions if something goes wrong or a critical operation is
earned out. In any case, it is recommendable to let one team design the basic system and
another team design the supervisory system to protect against making the same mistakes
twice. In this way it is more likely that all situations will be covered.

6.11 Opportunities for Spacecraft Commonality, Modularity
and Standardisation in Future Solar Science and
Applications Missions

During a luncheon speech, at ISU 96, on the international implications of smaller
cheaper, faster (SCF) spacecraft. Dr. Gregg Maryniak of the Space Studies Institute
hypothesised what SCF spacecraft might mean for the science fiction film industry, in
particular, for the script of the tenth or so Star Trek film [Maryniak, 1996]:

Chekhov: Captain! Sensors indicate three Starfleet Class M matched
handbags!

And Kirk will be cool...

Kirk: Steady Chekhov! Many bags look alike.

Although Maryniak was having fun at the expense of future "luggage" sized spacecraft,
the fact that his joke included more than one spacecraft (or handbag) and these spacecraft
(or handbags) looked alike, points in the direction of several important, but rarely

discussed, concepts behind SCF spacecraft. These three concepts are commonality,
modularity and standardisation, and they have practical benefits and potentially large
implications for future solar science and solar warning spacecraft missions.

Solar science spacecraft missions suffer from a lack of political, institutional and space
science community support which makes these missions a relatively low priority in some
space agency budgets. Chapter 6 has discussed several technological alternatives that
may make future solar missions less costly and more budgetarily viable. Chapter 3
discusses organisational solutions to gathering support for solar science research an
solar warning and forecasting applications. This section attempts to link the
technological solutions in chapter 6 to the organisational solutions in chapter 3 through
the system engineering concepts of commonality, modularity and standardisation. A
tentative plan will be introduced on how these three concepts can be used to foster cost
reductions in solar science and solar warning spacecraft through international co¬
operation.

6.11.1 Defining Commonality, Modularity and Standardisation

Before proceeding with a discussion of how the concepts of commonality, modularity
and standardisation can be applied to a high technology, international solar observation
framework, it is important to define these concepts. These three concepts will be
collectively referred to as SCM (not to be confused with SCF) throughout the rest of this
section.

6.11.1.1 Commonality

Commonality refers to the repeated use of the same component or system on more than
one spacecraft and is a measure of the versatility inherent in a single component or
system. Commonality is important in realising the economic and temporal benefits of
utilising SCM concepts in spacecraft design and relies heavily on standardised
requirements.

6.11.1.2 Modularity

Modularity defines the ability of a spacecraft design to integrate different components or
systems for different missions. Modularity can be thought of as the measure of the
universality of a spacecraft's interfaces and overall design. Modularity is enabled by
standardised interfaces, common components and systems, and clear reference designs.

6.11.1.3 Standardisation

Standardisation is simply the organisational task of setting and agreeing to abide by
defined component, system or spacecraft specifications for certain mission requirements.
Standardisation in the context of this section is especially critical for setting design
requirements, for building interfaces and for creating reference designs [section 6.11.1.4].

6.11.1.4 Reference Design

Another important term also used in this section is "reference design." A reference
design is a "blueprint" for a system or spacecraft that can be utilised as a generic and
adaptable baseline for further engineering to create a system or spacecraft design that
meets specific mission requirements. In the terms of this section, a good reference design
is a design that meets the needs of multiple users with minimal adaptation.

178 • Ra: The Sun for Science and Humanity

6.11.2 Rationales for Commonality, Modularity and Standardisation

SCM concepts, if successfully implemented, can create significant advantages in terms of
resources spent on spacecraft design and production and can thus decrease the cost of
solar observation. Additionally, several technological and political themes also serve as
rationales for SCM in future solar observation spacecraft design.

6.11.2.1 Technological Opportunities

Many currently emerging spacecraft technologies can leverage the operational flexibility
needed to create true SCM capabilities in solar probe and satellite designs. Pushing
technological limits too far can have detrimental effects on the ability of certain users to
afford, build and exploit an SCM reference design, but if properly combined and applied,
emerging technologies promise to make a single system or spacecraft design viable for a

wider, rather than narrower, group of users. The promising candidate technologies
include:

Non-chemical Propulsion Systems

Electric propulsion (solar and nuclear) and solar sail propulsion can endow a single solar
probe or satellite design with the capability to reach a variety of solar orbits or Lagrange

points.

High Density Power Systems

New power system technologies like lithium polymer batteries, gallium arsenide, indium
phosphide and multi-layer solar cells, and solar thermodynamic generators can increase
the total available power per unit mass of power system on a spacecraft over standard
batteries and silicon solar cells. Radioisotope generators (RTGs) also offer this capability
using proven technology. By incorporating larger power capabilities at less mass cost, a
single power system or spacecraft design can accommodate a greater variety of solar
instrument payloads and operational lifetime requirements.

Lightweight Alloy and Composite Structural Materials

If certain production challenges are overcome, lightweight alloys and composites can
contribute to solar probe or satellite structure mass reduction, which can contribute, in
turn, to the ability of a common spacecraft design to reach different orbits and Lagrange
points and use different launch systems.

Smart Structures

Adaptive systems and materials capable of reacting to external input rapidly, repeatedly
and autonomously through material properties or active electromotor input can allow a
spacecraft to adapt to different environments and vibration regimes.

Inflatable Structures

Externally and internally rigidized inflatable structures offer low mass and low cost
options for various deployable spacecraft components such as instrument booms and
reflector dishes.

Variable Thermal Systems

Microlouver, variable emissivity radiators are a promising technology capable of
enabling a single spacecraft design to operate in different temperature regimes.

Small, Lightweight Sensors

Military derived sensors can decrease the mass of the tracking system and instrument
payload for solar spacecraft while maintaining or increasing previous observational

capabilities.

Fibre Optic and Wireless On-Board Communication

Wire cables, cable harnesses and connectors occupy a noticeable mass fraction of any
spacecraft. The use of fibre optic cables or wireless communication (infrared beams,
radio signals or low power laser beams) on board a spacecraft can reduce the total mass
of a spacecraft introduce flexibility in data transmission.

isn • Ra- The Sun for Science and Humanity

Converging International Information Processing Standards

The increasing international standardisation and compatibility of computer hardware,
software and interfaces can contribute to the commonality of spacecraft information
systems.

6.11.2.2 Decreased Unit Development Costs and Time Frames

Once available technologies are correctly incorporated, an SCM reference design can
lower the development costs and time frame for the a new spacecraft. Instead of
"reinventing the wheel" for all of a given spacecraft's systems, those systems that are
non-specific or non-critical to the spacecraft's mission requirements can be lifted from the
reference design and applied to the new spacecraft.

6.11.2.3 Cost Reduction Through Economic Scales of Production

Utilising the same system or spacecraft for multiple missions will also reduce the
production costs of the system or spacecraft. Production costs are lowered because the
tools and knowledge needed to create one system or spacecraft do not have to be
modified to create an additional system or spacecraft. The learning curve that is

advanced by producing more than one system or spacecraft also contributes to lowered
costs.

6.11.2.4 Increased Scientific Return Per Unit Cost

With lowered development and production costs, the costs of scientific exploration are
also lower because more data can be obtained for the same investment.

6.11.2.5 Convergence of Science and Applications in Solar Observation

Proposed solar science observation missions hold many instrument and spacecraft
requirements in common with proposed solar warning and forecasting spacecraft. Solar
observation at various Lagrange points, solar stereoscopic observation, and instruments
for ionospheric and magnetospheric observation have the potential to satisfy scientific
curiosity and to provide data for improved solar forecasting models at the same time.
Project managers and engineers for solar science and solar warning spacecraft can
cooperate to design common instrument systems, support systems and spacecraft to
lower development and production costs.

6.11.2.6 Potential Synergistic Interaction of Cost Reduction and International Co¬
operation

International co-operation in space science projects and missions usually implies a higher
total cost for a particular project or mission because of the higher managerial costs
associated with complexity of international co-operation. In the past, international co¬
operation in space science has also been limited primarily to scientific data co-ordination
of independent agency projects and missions. However, by expanding space science co¬
ordination into the engineering of international projects and taking advantage of the
international demand for solar observation spacecraft, it may be possible to actually
reduce the costs of international co-operation in space science by designing and utilising
SCM spacecraft on an international scale.

Technology Challenges & Issues • 181

6.11.3 Trade-offs and Drawbacks to Spacecraft Commonality, Modularisation
and Standardisation

Designing for SCM in a system or spacecraft holds certain risks, and this section outlines
the risks that must be balanced against the benefits of SCM concepts described in section
6 . 11 . 2 .

6.11.3.1 Large Initial Development Costs and Time Frames

Although the development costs and time frames for future spacecraft that use common
systems are lowered, the cost and time frame needed to develop a common system that
can meet more than one set of mission requirements can be greater than designing the
equivalent system for one spacecraft.

6.11.3.2 Design Non-optimization

Even a very flexible SCM spacecraft design will not meet the requirements of every
potential user. Unique but critical requirements must be addressed by a separate
spacecraft or by a modular system that can interface with the basic SCM reference design.
Although an SCM design may meet the requirements of a variety of users, it may not
meet them all in an efficient manner. A minimum of overdesign in certain system
capabilities will be needed to make a design suited to the mission requirements of more
than one user.

6.11.3.3 Potentially Limited User Demand

Care must be taken when defining potential users for an SCM spacecraft and obtaining
development funding from them. Commitments from multiple groups to utilise an SCM
spacecraft may be needed before the additional funding and development necessary for
SCM can be undertaken. If an SCM design does not meet the needs of more than one of
its intended users, the additional funds needed to design for SCM are wasted. If multiple
user demand is not viewed as likely early in the design process, SCM concepts should not
drive that process. If enough users are found to warrant SCM, it is critical to build to
those user needs (possibly with some negotiation between different user needs)
throughout the design process.

6.11.4 A Short Synopsis of Spacecraft Commonality, Modularisation and
Standardisation in Space Science: The Tale of Two SCM
Programmes

SCM has long been a goal of spacecraft designers since the earliest satellites were
launched. Communication satellite families achieved SCM early in their development,
and some commercial, military and civil government remote sensing satellites are
currently converging on SCM designs. Science satellites and probes, however,
experienced a less successful advance towards SCM over the same time period. This is
partly because of the unique requirements that science missions impose on spacecraft
payloads and buses through their different observation objectives and their varied
operating environments. These requirements simply made SCM impossible or very
costly using past, mission specific technologies. Many of the emerging technologies
described in section 6.11.2.1, however, are not specific to any particular mission; rather,
they increase the flexibility of a spacecraft or system by increasing its support and
performance capabilities. The lack of SCM concepts in science spacecraft design is also
attributable to the dual goal orientation of most space agencies throughout the world
which teams scientific exploration with technology development in the same
programmes. NASA has taken steps to remedy this situation through the separation of
scientific missions in its Discovery programme from technology development missions in

182 • Ra: The Sun for Science and Humanity

Its New Millennium programme. This new technological and programmatic
environment provides an opportunity for SCM to be achieved and applied in various
spacecraft missions, including solar observation. Before describing how SCM might
specifically benefit solar observation in a stepwise progression, it is important to contrast

two purposeful efforts, one past and one present, towards SCM in science spacecraft
design. r

6.11.4.1 Goddard Space Flight Center's Multimission Modular Spacecraft (MMS)

In the early 1970s, NASA's Goddard Space Flight Center recognised the need to develop
a large, adaptable spacecraft bus to support future orbital observatories. To capture the
most astrophysics and Earth sensing missions in one spacecraft, the MMS focused on four
missions: solar. Earth and stellar observation from LEO, and Earth observation from
GEO. The MMS bus incorporated only power, attitude and control, command and data
handling and thermal systems on a triangular, prism-shaped support structure.
Instrument payload, additional solar power and propulsion were all mission specific and
integrated on the top and bottom of the support structure via transition adapters. MMS
was compatible with the Delta, Atlas, Titan and Space Shuttle launch. [Falkenhayn, 1987]

MMS followed several design rules to obtain its CMS capabilities: one thermal design for
all missions, maximise the use of qualified and standard NASA components, minimise
electrical and mechanical connections at interfaces, and no thermal break at interfaces.
Testing and competitive procurement was placed at the system level to guarantee
modularity. The MMS created cost advantages in total spacecraft design by reducing
spacecraft integration and test time. MMS held interfaces standard but permitted

modular system upgrades to improve performance and create design flexibilitv
[Falkenhayn, 1987] y '

In summary, MMS achieved limited SCM advantages with proven technology by
designing a common service bus with modular components that could interface with
different propulsion systems and instruments payloads to accommodate different
mission requirements in a common environment. MMS reduced costs and development

time frames by applying SCM concepts to users with common support system
requirements. J

6.11.4.2 Jet Propulsion Laboratory's New Millennium Programme

hi contrast to MMS, the Jet Propulsion Laboratory's approach to SCM in its New
Millennium Programme (NMP) is driven more by technologies that enable SCM than by
meeting the common needs of several users. One of NMP's Integrated Product
Development Teams is dedicated to Modular and Multifunctional Systems (MAMS)
Instead of designing a standard service bus with modular systems and common
interfaces, MAMS is concentrating on exploiting technologies to combine multiple
functions (propulsion, power, structures, mechanisms, thermal systems) into single
systems. One of the best examples of the MAMS approach is an inflatable reflecting dish
that can be adapted for long baseline interferometry, in subsurface planetary sounding, in
remote sensing radar, in soil moisture radiometry, for a submillimeter space telescope
and as a space power antenna. Another example of a MAMS concept is a
micropropulsion unit for miniprobe propulsion or precision station-keeping in larger
spacecraft. MAMS drives SCF through multifunctional SCM systems that significantly
reduce overall spacecraft mass and enable open spacecraft architectures. [NMP Events
Theme 10 Homepage, WWW]

6.11.5 Future Opportunities to Incorporate and Exploit SCM Concepts in
Solar Observation Spacecraft Design

Future solar observation spacecraft for solar science and solar warning systems have
opportunities available to them to take advantage of both the MMS and MAMS
approaches to achieving SCM benefits. These opportunities stretch across the near-term,
mid-term and far-term Ra Strategic Framework.

6.11.5.1 Cluster Phoenix: A Near-Term Opportunity for International
Commonality and Standardisation in Solar Science

The loss of the Cluster constellation presents ESA and possibly other space agencies
involved in the International Solar Terrestrial Physics Programme (ISTP) with the
opportunity to apply SCM concepts immediately and at low investment to replace
Cluster's capabilities. Although ESA management is currently leaning towards
launching the Cluster spare satellite as soon as possible to complement ISTP data in the
magnetospheric cusp region, ESA should also consider not wasting its Cluster
development investment and procure three more common Cluster satellites for a future
launch. Alternatively, if Cluster procurement funds are not available, ESA should look
outside its programme for a small satellite design that can carry the most important
Cluster instruments to complement the Cluster spare satellite. Possible candidates might
include university minisatellites [section 6.11.5.2] or a proposed NASA second generation
space physics and particles microspacecraft [Second Generation Microspacecraft
Homepage, WWW]. The Cluster loss could provide an international driver for ESA,
NASA and other space agencies to advance independent, but coherently related,
development of small, common, standardised solar observation spacecraft in co¬
operating countries.

6.11.5.2 University Microsatellites, Military Minisatellites and Commercial Buses:

Mid-Term Opportunities to Exploit SCM Concepts for Solar Science and

Solar Warning Spacecraft

In the mid-term, space agencies involved in the ISTP programme should take advantage
of existing and developing modular university microsatellites. For example, Surrey
Satellite Technology Limited, a company formed by the University of Surrey in Great
Britain in 1985, currently offers the Micro-Bus, a modular microsatellite platform that
houses systems and payloads in customisable tray modules [Micro-Bus-SSTL Modular
Microsatellite Platform Homepage, WWW]. A Micro-Bus satellite can be developed in as
quickly as 9 months and offers university and agency researchers involved in ISTP the
opportunity to quickly and inexpensively obtain additional data about a particular
phenomenon when the current ISTP constellation and instruments prove to be
inadequate. University minisatellites can also be utilised for technology demonstration,
especially the flight validation of new, lightweight sensor technologies for future solar
missions. Stanford University in the United States has developed two SQUIRT (Satellite
QUIck Research Testbed) microsatellites, one of which is known as SAPPHIRE (Stanford
Audio Phonic Photographic Infrared Experiment). SAPPHIRE is flight testing a
micromachined infrared sensor for NASA's Jet Propulsion Laboratory [SQUIRT
SAPPHIRE Homepage, WWW].

Some SCM technologies and platforms previously developed by the U.S. Department of
Defence for its Strategic Defence Initiative and by its Ballistic Missile Defence
Organisation may also be applicable to solar science or solar warning spacecraft. The
U.S. military is currently developing Clementine II, a miniprobe bus nearly identical to
the now famous Clementine I spacecraft, which is capable of launching, monitoring and
controlling three identical, high thrust, daughter spacecraft designed for asteroid
interception [Worden, 1996]. The Clementine bus and daughter spacecraft might be

184 • Ra: The Sun for Science and Humanity

easily adapted to the deployment of a solar sensor constellation in a libration orbit
around a Lagrange point. Stanford University is also pursuing a mother microsatellite
capable of launching four identical daughter picosatellites through its second SQUIRT

SSriu (< ?. rbiting Picosat Automatic Launcher) [SQUIRT OPAL Homepage,

yvWWJ. The U.S. military has also developed a modular minisatellite structure design
for its own sensor demonstration needs. Known as MSTI (Miniature Sensor Technology
Integration), the third spacecraft in this series has been adapted to track warm objects in
space but its ability to gather background clutter data has the ability to derive data on the
solar interaction with the Earth's atmospheric limb, solar scattering effects and solar
specular intensities [Barnhart, et al, 1995]. Future MSTI spacecraft might be guided
towards more direct solar phenomena observation missions.

Modular commercial satellite buses may also prove to be adaptable to certain solar
observation missions. Lockheed Martin recently offered its LM700 bus which can
accommodate distributed payload components or more modular payloads up to 500 Ibm.
The LM700 uses a graphite epoxy structure to reduce weight, features gallium arsenide
solar cells and can launch on Proton, Delta, Long March and LMLV-1 vehicles [LM700
WWW], Although designed for remote sensing and surveillance missions,
the LM700 can attain two nadir orientations and has two-axis gimbals for its solar arrays
which could permit certain solar observations. Although not ideally suited to solar
science, the LM700 might prove to be a cheap means of creating a dedicated solar
warning and forecast satellite. Alternatively, the modular university or military micro-
and minisatellites described above could be used to create small solar warning
observation networks. 6

By exploiting existing SCM and SCF spacecraft in academia, industry and the military
cheap, quick response solar observation missions could be mounted in the mid-term to
support current and planned solar science efforts (ISTP and FIRE). These existing
spacecraft might also provide the first dedicated platforms for space-based solar warning
and forecasting instruments in Earth orbit or at various Lagrange points. Use of these
spacecraft will also be crucial in flight testing instruments and gaining experience in SCM
design concepts for a new generation of in situ solar observation spacecraft.

6.11.5.3 An International Reference Bus Design for Solar Observation: A Far-Term

Opportunity to Pursue SCM Concepts to Sustain Multiple, Long Duration, In

Situ Solar Missions

The Ra Strategic Framework realises the scientific need for dedicated constellations or
networks of solar observation platforms beyond the Earth orbit and Langrange point
spacecraft discussed thus far. Such spacecraft may also prove crucial to extending solar
warning lead times and improving the accuracy of solar forecasting beyond the
capabilities envisioned even with dedicated, Lagrange point spacecraft. Although the
Framework predicts that these spacecraft will occupy different solar orbits (polar
synchronous, etc., see section 10.1) and will carry different instruments (stereoscopic,'
neutral atom imagers, etc.), the environments in which these spacecraft will fly and their
possible payloads do not impose radically different or impossible design requirements
especially when the technologies of section 6.11.2.1 are taken into consideration. In light
of their common, baseline requirements, it is recommended that the international
community pursue the design of a standard, common bus for in situ solar observation
constellations and networks. This bus should be a reference design only, adaptable to the
needs of several solar observation missions, but not contingent on planned national space
agency or solar warning and forecasting missions. The bus design, its requirements, its
standards, and its interfaces would be hashed out through an international forum similar
to various international scientific working groups but endowed with an engineering
emphasis. Section 3.2 presents the organisation of a proposed international solar working

Technology Challenges & Issues • 185

group, which includes an engineering section dedicated to the creation of a solar
observation service bus reference design. Once the reference design is available, national
space agencies can utilise it as a baseline to save development costs and time frames by
adapting it to their specific solar observation mission needs through their own modular
payloads. The commonality of the service bus reference design would allow space
agencies and solar warning and forecasting organisations to pursue independent projects
while co-ordinating to engineering costs and time frames. By involving the international
community, the demand needed to justify an SCM reference design for solar observation
spacecraft networks and constellations is met, and its benefits distributed to the
maximum number of solar observers.

186 • Ra: The Sun for Science and Humanity

Chapter 7

Market and Funding Issues

In this chapter the market issues and the possibilities of funding for the Ra Far-Term
programme are discussed. When we say market, we refer to the interaction between the
parties in a given business situation. The involved parties are the scientific community,
the public and private sector, private industry, education and entertainment. In the first
section, we will discuss the market and its related issues, in the second the funding
sources, means and methods and in the third the marketing.

7.1 Markets for Ra

In the search for potential and existing markets for Ra, the following ideas have been put
forward. Up to now, the results from missions performing solar measurements and
acquisition of relevant data, are mainly used by the scientific community and the space
environment prediction services. It is important to differentiate between profitable
markets and non-profitable markets. The non-profitable market in the case of space
environment prediction is made up of elements in the public sector that exists more or
less as a public good. They distribute the current space environment predictions at no
cost. Does this mean that there is no profitable market for space environment prediction 7
Absolutely not! You can always charge money, if your product is of value to the
customer. We have found an example where power companies pay for research and
customer adapted space environment predictions [Lundstedt, 1996]. And this is done
even though the power companies can get predictions at zero cost. There is an added
value for the product! Another reason for not relying on the institutes giving predictions
is that they do not have a responsibility to provide predictions during crisis such as wars
You add value to the service/product by providing more reliable predictions, longer alert
time etc. The conclusion is that there is a profitable market for space environment
prediction. You can even create a market through the development and provision of
customer tailored products, in this case customer adapted prediction.

7.1.1 Space Environment Prediction

As part of the market survey and evaluation, the current end to end chain of users of
solar data for space environment prediction was examined in terms of interest,
opportunities, opportunity costs and market growth potential.

Fig. 7.1 End-to-end chain of users of space environment prediction.

Space environment prediction services get input data for their models either from
institutes or space agencies. The prediction is done and then delivered to the customer as
schematised in figure 7.1. The customer might be a space agency, a power company, a
satellite operator or insurer. The big questions are : "Is there an end user willing to pay
for the service?", "How big is the market?", "How do you estimate the size of the
market?" and "What is its growth potential?"

One way of estimating the size of the market is to ask industry how interested they are in
paying -for the specific service. This proves to be quite difficult, because the companies |
cannot estimate the value of a service until they see the direct benefit of the
service/product. However, we have obtained from an electric power company the cost
of a lightning locating system as being $20,000 U.S. annually [Andersson, 1996]. If a
commercially available space environment prediction system would exist, this cost could
be seen as a maximum [Andersson, 1996].

Another approach would be to see how much it costs for companies not to use the
service/product. How much does it cost when a telecommunications satellite is
destroyed by magnetic storms or high energy particles? A lot of research has to be done
on this point. How much does it cost when an electric power net goes down in Canada as
a cause of magnetic storms? The costs of the power breakdown of the Hydro Quebec is
estimated to be more than $10 million U.S., but much higher costs have been estimated
for consequences of the breakdown. This is also a reason for insurance companies to look
into this matter. Good space environment prediction could prevent many expenses for
the insurance companies.

The real issue is the pressure that a customer exerts, for example on a power company. If
it is vital to have power continuously, the customers simply say they are prepared to pay
for the extra service. The service in this case is in the form of space environment
prediction used by the power company to deliver a more continuous service to the

188 • Ra: The Sun for Science and Humanity

customer. So the push towards the use of space environment prediction starts at the
customer. The customer could be a bank or financial institution that needs 24 hours
continuous information services or a hospital that needs continuous power.

How much is the cost of a lost life in a remote part of Australia (the country highly relies

on radio linkss ) as a cause of bad radio communications during a magnetic storm
[Thompson, 1996]?

The total annual space environment prediction market is about $100 million U.S. at the
moment [Worden, 1996]. It is expected to increase up to $200 million U.S., within the
next ten years.

There is a clear demand for continuous space environment prediction which is more
precise (at the moment 30-50%) and has a longer warning time [Worden, 1996].

There are indications that changes of the space environment have an influence on the
Earth weather, and even, in some circumstance, possibly our human health [Atkov 19961
[Campbell, 1996]. ' J

Let's have a look at the market of end users for prediction. Two interesting future end
users are the electric power industry and the planned satellite constellations for mobile
communication. The satellite constellations are made up of large numbers of satellites,
some in low Earth orbit and some in medium Earth orbit. The estimated total budget of
these nets is somewhere between $10 and $25 billion U.S. How much are they prepared to
pay for space environment prediction? The answer depends upon other things, such as
the quality of the prediction and what countermeasures can be applied during high solar
activity. It is mostly the upper constellations that are interested in space environment
prediction.

The efficiency of the commercialisation decreases as the technology matures. Space
environment prediction is still an immature product and therefore interesting. Do not

miss the window of opportunity. Space environment prediction has its window now 1
Use it!

7.1.2 Entertainment and Education Market

By converting the scientific results, partly and appropriately, into entertainment, two
results can be obtained:

1. Increased public awareness and increased interest for solar science;

2. From the entertainment market the Ra scientific missions can be partly
funded, if Ra shows the market potential for entertainment.

The entertainment market is big and even a small part of the market can generate large
sums of money. However, the market is very sensitive to market pressures [section 7.3].

The prediction of auroras is an example of a combination of entertainment and education.
The recording and telecasting of such solar generated phenomena can be a core element
of televised documentaries. Taking spectators up in a helicopter to view the aurora at the
right moment provides another business opportunity.

Market and Funding Issues • 189

Another example of a potential entertainment market is the "Las Vegas mission", referred
to as the suicide probe in section 10.2. This spacecraft is launched, toward the Sun, from
a mother spacecraft. It will obtain valuable scientific knowledge about the Sun.
Moreover, it is very special that a human built spacecraft will reach the Sun at such a
close distance.

The "Las Vegas mission" can be a source of a lot of entertainment. Like, big gambling
events. How long will it survive? Is it still alive? The name Las Vegas is strongly
connected to gambling through the town in U.S. A suicide probe to the Sun will be
consumed by the Sun. The big question is when? No matter how good all the
calculations and estimations of the lifetime of the probe are, no one will know for sure
how long the probe will survive until it actually is consumed by the Sun and its
environment. This gives an excellent opportunity for gambling. Can you see the
headlines "How long will Vegas make it?" or "Latest update from Vegas, temperature
has now reached 600K and is rapidly increasing" in combination with pictures of coronal
mass ejections.

7.1.4 Science Market

More and more contracts between universities and industry are being made. This is a
way of getting funding for science. Some of the scientific questions are [section 5.1]:

What are the causes of coronal heating and coronal holes?

What are the causes of CMEs?

What is the origin and acceleration processes of the solar wind?

How different is the polar solar wind from the equatorial?

Does any change in the Sun also effect change in Earth weather/climate?

What causes the solar constant to change?

The universities or institutes perform research that is relevant to the industry and thereby
receive fees, funds and/or grants. In this lies a big potential. In the case of space
environment prediction this could be very interesting to power companies for example.
Why not leave the leading role to industry as part of their Research and Development.

7.1.5 Expected Time Evolution of the Markets

To predict the evolution of the described markets is highly speculative. However we
envisage a combination of the following factors [figure 7.2]:

• Science is more and more related to direct application of industry,
therefore, it is expected that the space environment prediction market will
increase [Worden, 1996] and the science market will be stable.

• To increase the funding for science, the public awareness concerning
science has to be increased. A possible way to do this, is to increase the
entertainment related to science or to increase benefits from scientific and
efficient technological knowledge to develop and implement light and
heating infrastructures for buildings and transportation. This results in
potentially high and significant reductions in energy costs and significant
influence on the health of the global population [section 10.4],

190 • Ra: The Sun for Science and Humanity

7.2 Project Funding

The sources of funding for Ra may be divided into three major parts. The first one is
governmental funding, the second is private funding and the last is a combination of
them both. The funding can also be spread along a time-scale.

7.2.1 Governmental Funding

Governmental funding can be civil, military, agency, institutional funding etc. It can be
from a single source or from a combination depending on the specific project, its
characteristics and national and/or industrial interests [section 3.2], Funding decisions
for Ra can also be made by organisations led by national politics. °

The borderline between military and civil funding is not always clear. This is a case for
dual use technologies, where a project might be of use for both civil and military

purposes. In some space agencies the difference is clear. ESA only funds non-militarv
projects. J

In some places on Earth, radio communications that are influenced by the space
environment can mean the difference between life and death. In other places the space
environment affects public power networks. There is also military interest in the space
environment. The list could be long and the intention is just to show that there is
governmental interest in predicting the space environment.

Governments tend to have a shorter and shorter perspective in the sense that they have a
higher priority in the near-term. They want to see a quick return on the investment for a
public service in order to enhance their political strength.

Getting public interest in the Ra programme is likely to increase the availability of
governmental funding sources [chapter 3]. We think that further studies on this should
be done.

One of the big fund-raisers is the scientific community, irrespective of national
boundaries. For science there is a governmental interest since science can create of
benefits to the society. Scientists tend to be good in raising money from governments,
funds, institutes and industry. But why not increase modest amounts of their funding
effort. Let the scientists think and act commercial. Let scientists put on a fancy dress and
talk to people! Do research that is relevant to industry and make contracts with them.
Power companies could, and some already do, pay for science on space environment
prediction [Lundstedt, 1996]. Go over the blocks in the funding? The Sun is the biggest
plasma laboratory we know of. Why don't space physicists and nuclear physicists co¬
operate more? There is co-operation between , but it could definitely increase. There
might be a problem with the two separated budgets (e.g. that is the case in U.S.), but with
a bit of goodwill and enthusiasm they should be able to overcome such problems and
thereby increase the total funding for Ra.

Another supra-national global funding source is the United Nations (U.N.). Lives of
people all over the world can be saved and the life quality can be increased by using
space environment prediction [section 3.4]. This is in the interest of U.N. and of all
humanity.

Space agencies are interested in solar science and space environment prediction. This is
extremely important for the manned programmes. Improved measurements and models
should benefit the manned space programmes and thereby constitute a ground for
funding.

7.2.2 Private Funding

Private industry is an alternative funding source for some elements of the Ra project. It is
easier to see the direct benefits of applications for industry rather than the benefits of
science on the Sun and its effects on Earth (even though there are benefits from science).
Typical applications for industry could be space environment prediction. Potentially
interested private parties in this domain are communication satellite operators and
electric power companies in some countries. These two industries are big and they
sometimes need better predictions than a "general" space environment prediction
institute offers. Electric power distribution companies are large infrastructure companies.
Satellite communications are increasing rapidly. A number of different satellite
constellations are planned for mobile communications. Some of them will use satellites
in low Earth orbit and some of them will use satellites in higher orbits. The average
budget for each constellation is some 3 billion US dollars [Pelton, 1996]. The launching of
the satellites in these constellations starts 1998/1999 and the volume is well over one
hundred satellites. Space utilisation is very expensive and it will be affected by the Sun
and its environment, therefore funding for Ra should take a prominent place in their
priorities. Is there enough flexibility in their business plan to pay for the service?

There seems to be a lack of awareness in private industry about the Sun and its influences
on Earth. If this situation could be improved it would surely be easier to find funding for
Ra. Another problem is that private industry still knows that space is risky (high
insurance premiums). This makes private funding more difficult. There is a big need of
finding ways to show private industry that with proper insurance and technical measures
space business does not have to be any riskier than any other industry.

One way to get funding money or risk money is to use small companies outside of the
space field who wish to enhance their image through space work. You could argue that it
introduces more risk, but that remains to be proven.

IQ? • Ra: The Sun for Science and Humanity

7.2.3 Combination of Private and Governmental Funding

If you can show technical and financial feasibility and if a market can be determined,
private money funding could be invested together with governmental funding. "In
today's environment shared funding is a prerequisite to get things going" [Cohendet,
1996]. One way of having combined funding is to let private industry build, finance]
operate and transfer the project to the government. This is called concession funding.
One of the difficulties here is who takes the risk. An alternative to this is to do it the other
way around. This is motivated by the fact that private industry tends to be a more
efficient operator.

To make this type of funding possible industry must show some interest in the Ra. Any
will to invest, even if it is a small investment, is enough to show the space agency that the
industry is interested. Space agencies on the other hand, should encourage non¬
aerospace companies to invest in Ra.

There seems to be a lack of interaction between potential users and sellers of solar data
and applications. Improving the interactions between government and private industry
would facilitate and improve the funding opportunities among industry. We think that
there is a lot to improve in these areas and further studies on this should be done.

In some situations clusters of companies are very competitive. Could a cluster consist of
electric power companies and space environment prediction institutes/companies? The
answer is, yes it could, it already exists in Sweden [Lundstedt, 1996], And this can
expand to a global scale. In the U.S., similar suggestions have been made to let power
companies invest together in geomagnetic storm prediction, but so far nothing has been
done on that point [Worden, 1996]. You have to have a strong force or personality acting
on the decision makers. In Canada the power companies use several different space
environment prediction resources. The mere fact that the power companies have shown
an interest in our investigations is significant. There are other clusters that could be
interesting, e.g. communication satellite operators and space environment prediction
services. The interesting thing is that clusters often have a competitive advantage. Are
they willing to fund Ra? It depends on the market situation. Furthermore the clusters
serve as development centres with strengthened competence and feedback. You get a
situation where end users are innovators.

There is a trend toward letting contracts between universities and industry. The
universities do research relevant to industry and the industry funds part of it. This is also
a way of getting combined funding. We also see a trend where solar activities are
moving from research driven to product/service driven.

7.3 Marketing

Space businesses have a lot to learn from private industry concerning marketing
diversification and the creation of new markets. In general the market demand is a
function of the marketing effort as seen in figure 7.3. By increasing the marketing effort,
the market for solar data can be increased. J

Among relevant aspects for Ra, are the importance of positioning the product on the
market, in the correct market segment and in the customers requirements' domain, to
convince the future customers why they need the use of space environment prediction]

Market and Fundine Issuer • 1 W

For the space environment prediction, a way to increase the market is:

• show that a lot of satellite losses are due to magnetic storms/high energy
particles;

• show power companies that they have increased power consumption in the
transmission lines because of the magnetic storms;

• quantify the losses, to obtain a profit-loss calculation.

Fig. 7.3 Market demand as a function of marketing effort.

The marketing of the entertainment is based on perception and less based on rational
thinking. Some examples of marketing are:

• Use famous persons to talk about solar physics and the space environment.
Television personalities are examples of people that attract other people to
listen. And why not? You do not have to follow the traditional way of
doing things;

• Use solar relevant entertainment. Virtual reality trips to the Sun or a
stereoscopic view of the corona;

• Use solar science in the public education. This gives a broader interest
understanding for solar science;

• Make television programmes and contests related to the Sun for children.
Contests have a multiplying effect [Willekens, 1996]. You only have one
prize but a lot of people in the contest and a lot of viewers. As an example:
ESA had a "space theme" at Disneyland Paris.

194 • Ra: The Sun for Science and Humanity

Chapter 8

Near-Term Programme

8.1 Overview

This chapter will provide the details of the Ra Near-Term Programme as introduced in
chapter 2. As described in that chapter, "near-term" is from now until the year 2000.

Each part of the programme described in the following eight sections is relatively low in
cost and either builds on existing systems and infrastructure or requires only modest
developments. We believe the recommendations are realistic and play an important role
in realising the objectives described in chapter 5. They also provide a foundation for the
programmes described in the mid- and Far-Term Programmes. To build on existing solar
observation instruments (namely SOHO) and to continue with a logical sequence of solar
observation satellites, we discuss the Cluster replacement programme [section 8.2], As
we believe space environmental forecasting will become more important to the space
community in the mid- and far-term, we recommend immediate work on improving
forecasting models [section 8.3]. As the amount of archived data continues to grow and
additional solar observation satellites are launched it becomes ever more crucial to ensure
the co-ordination and accessibility of both the new data and those from the past
[section 8.4]. Then, in section 8.5, we describe the near-term implications of the Working
Group for International Solar Exploration & Application (WG ISEA) [chapter 3], To help
advance the mid- and Far-Term Programmes through to fruition, we envisage increasing
awareness of solar science and solar terrestrial connection, thereby fostering support
beyond the scientific community [section 8.6]. The Near-Term Programme is concluded
with reasons to support the "faster, cheaper, and better" concept into future technology
development.

8.2 Replace Cluster

The four original Cluster satellites were lost on June 4th, 1996 with Ariane 5's maiden
flight failure. They, together with the Solar Heliospheric Observatory (SOHO), were to

be part of ESA's Solar-Terrestrial Science Programme (STSP), and part of the International
Solar-Terrestrial Physics (ISTP) programme. The timeline of a Cluster recovery is
governed by the desire to achieve simultaneous observations with other ISTP Spacecraft
[Cluster within STSP]. ISTP includes STSP and spacecraft from the United States, Japan,
and Russia, and aims to investigate solar-terrestrial physics and the Earth's
magnetosphere.

Hence the loss of Cluster has not only destroyed that mission but also deprived both
programmes of extensive valuable data, making the issue of replacement a critical one
among the international scientific community. For example, NASA Office of Space
Science (OSS) "roadmap", which develops a strategic plan for future space science
missions, relies partially on Cluster in its near-term plan [NASA's roadmap, WWW].

8.2.1 ESA Science Programme Committee's Work on Cluster replacement

The replacement of Cluster is currently being studied at ESA, and its implementation has
indeed already begun. "Everybody agrees with the principle that we should at least
partially recover the Cluster mission" quoted from Balsiger in Space News [de Sedling,
1996]. At the 3 July 1996 meeting in London, the Science Programme Committee
approved the funding of ECU 30 million to build the flight spare spacecraft of the first
Cluster mission, called Phoenix, and have it ready to launch by spring 1997. A decision
on a comprehensive replacement strategy is planned for November 1996, and four
options are being considered so far :

1. Fly Phoenix as soon as possible, which means maybe on 502 or 503 Ariane 5
launch, and build nothing else.

2. Fly Phoenix as soon as possible, and build 3 new Cluster spacecraft to go up
later for an estimated additional cost of ECU 350 million. 3 to 4 years are
required for the construction, which enable a launch by 2000/2001. At that time,
SOHO and the ISTP fleet will probably still be operational. Note that, along
with the 3 new Clusters, ESA will have to build another flight spare.

3. Hold Phoenix, build 3 new Cluster spacecraft and launch them together.

4. Hold Phoenix, build 3 national mini-satellites to accompany it, and launch them
together.

As of today, the second option has gained the most political support, for the following
reasons:

• We need to get the unique Cluster instruments, even just one set, into the unique
magnetospheric-cusp-region orbit, and contribute to the ISTP fleet as soon as
possible. There are 10 ISTP spacecraft in orbit, and Cluster contributes a lot to
the synergy.

• The cost of building three new spacecraft is easily identifiable and quite low,
because no R&D is necessary, but who knows the cost and politics of building 3
small new satellites? What instruments would be jettisoned for example?

• We really need four spacecraft to do the subtle 3-dimensional gradient
measurements needed in the solar wind.

• The flight spare as well as the new ones will be built by Dornier, basically the
same people will do exactly the same things as before to keep costs down.

196 • Ra: The Sun for Science and Humanity

8.2.2 Ra's Recommendations

The Ra Strategic Framework strongly supports a Cluster recovery mission. The question
is what form the replacement spacecraft should take. ESA should certainly explore the
possibility of using new technologies to reduce cost while still retaining capability. As
well as helping Cluster, this would also improve technology development for future
space physics missions such as applications-oriented solar-terrestrial monitoring
constellations.

Assuming an early launch of the original flight spare, a second important point is that by
the time the replacement spacecraft are launched (maybe 2001), the old flight spare may
well have ceased operating. Hence, provided adequate science instrumentation is flown,
the building of four Cluster replacements would seem prudent to guarantee the scientific
viability of the recovery mission.

Fig. 8.1 A low-cost alternative for Cluster recovery?

8.3 Improve Forecasting Models

The current state of affairs of the space environmental forecasting community has been
compared with the state of terrestrial weather forecasting over fifty years ago [National
Space Weather Program, 1995]. While there is a definite need for more measurements to
provide better forecasting capabilities, spacecraft sensors alone will not perfect the
forecasting job [Zwickl, 1996]. New measurements will need better forecasting models to
exploit the new data. °

8.3.1 Observations of Today's Space Environmental Modelling

The U.S. National Space Weather Program Strategic Plan, completed in August 1995,
outlined specific recommendations for space environmental forecasting. The authors'
recommendations for modelling included replacing the existing models with physics-
based quantitative models, transferring research models into tailored operational ones.

integrating models, evaluating them and making future models easy to upgrade
[National Space Weather Program, 1995]. We wholeheartedly agree, but we go beyond
those American recommendations to extend the concept internationally.

Current systems used in space environmental forecasting organisations are mainly
climatological and parameter-driven and many are quite old. For instance, many of the
forecasting models that the United States Department of Defense's 50th Weather
Squadron uses to predict ionospheric radio wave and solar event propagation were
developed between 1976 and 1982 [Lindsey, 1996]. Recent efforts have, however, been
made to acquire new specification models such as the Magnetospheric Specification
Model (MSM) developed at Rice University. These specification models are now in the
process of being converted to forecasting models.

Research efforts to predict and characterise the space environment have been on-going
for several years. A quick search of the Internet will yield many space environmental
models developed from a variety of places. Some research efforts have been made to
characterise solar flare propagation and model the interplanetary magnetic field
[IZMEM : IZMIRAN Electro-Dynamic Model, University of Michigan, WWW]. However,
the 50th Weather Squadron, for instance does not have any forecasting models for either
of these [Scro, 1996]. Other research projects show great potential for transition into the
space forecasting community to replace current systems such as the Lund Space Weather
Model which makes use of a neural network to predict a geomagnetic storm index [The
Lund Space Weather Programme, Lund University, WWW]. Operational benefits from
this and other research efforts, however, have not yet been realised [National Space
Weather Program, 1995].

Another problem today with space forecasting models is a lack of co-ordination between
the models. Currently, models are run independently of each other and do not provide a
cohesive picture. Future forecasting and specification models must include feedback
loops to "couple" the models. Coupled models are necessary to provide a clear picture
from the Sun to the Earth for the space forecaster [Scro, 1996].

8.3.2 Acquisition of New Models

Clearly, some work needs to be done with operational space environment forecasting
models. First, we recommend that a comprehensive international study be performed to
compare the effectiveness of current space environmental forecasting and specification
models. While the NSWP calls for verification of new models, there is no independent
agency today tasked with validating even the existing ones [Lindsey, 1996]. This study
would provide a baseline determination of which space environmental forecasts are good
and which ones need more work and will provide a mechanism for validating proposed
models.

Currently, very little money is budgeted for acquisition of new models. The Space
Environmental Centre, for instance, has personnel that develop new models and try to
keep apprised of research models that may be of use to the operational community
[Detman, 1996]. We recommend a new approach for model acquisition.

A New Approach to Model Acquisition

We envision a suite of co-ordinated industrial contracts be competed in the appropriate
countries with consortia of universities for acquiring new space environment forecasting
models.

The mode s competed for should provide data from the Sun to the Earth and be coupled,
his will mean, for instance, that an electric current predicted in the magnetospheric
model will be used as an input into the ionospheric model. The models should also
employ first-order analytical methods as much as possible. Empirical modelling should
be used where the physics is not well understood. The system should be easy to upgrade
for incorporating new knowledge and using new measurements. Finally, some sort of
neural network or other form of artificial intelligence will be needed to fuse the models
into a cohesive unit that will provide meaningful forecasting information.

The consortium should consist of universities that represent the fields of study in the space
environment. Universities that specialise in solar phenomena, magnetic fields, plasma
propagation, the ionosphere, the magnetosphere along with modelling specialists should
work together to develop the new models. We believe that models developed at the

university level as opposed to commercially derived models will be the most cost-
effective.

The industrial contracts should be independently planned but placed on an internationally
co-ordinated milestone schedule and modestly funded initially by agencies that will want
to use the models. These users would include the space forecasting and scientific
communities. Future versions of the models will of course be more expensive and will
provide increased accuracy. Requirements for the models should be established by the
users, and in the case of space forecasting users, the customers of the users. Future
models must provide the precise forecasting information that the affected customer

8.3.3 Summary of Modelling Recommendations

• Perform a correlation study to determine the reliability of current forecasting
and specification models so as to determine areas for improvement.

• Acquire new coupled, physics-based models that are easy to update by use of
internationally independent but, co-ordinated university consortium industrial
contracts.

• Derive requirements of new operational models through interaction with
proposed users and affected customers.

8.4 Co-ordinate and Apply Science Data

From section 2.2.4 "Past, Current, and Planned Missions" we know there already exists a
large amount of data related to solar activity. From chapter 5 "Objectives and
Requirements" we know there is a wide range of science and application objectives. Any
future direction in solar data observations should consider not only what data have been
collected but also how those data have been analysed, for what purposes, and how they
may be usefully integrated into future work in various fields. In this section we will
describe the impetus for co-ordinating solar data and then discuss some possible means
to achieve co-ordination.

8.4.1 Need and Opportunities for Co-ordinating Solar Data

It is interesting to note a conclusion made in 1970 (!) that

r ) ow be i n g achieved in measurements of electron and proton
distTibution functions is remarkable, and indeed is beginning to strain our ability
to absorb and comprehend the data [Manno and Page;i 970 ] y

Near-Term ProPTammp • iqq

This quote highlights the old idea that information and data are not enough but that
meaningful work requires comprehension. Also, consider the following :

The SOHO observations, in conjunction with co-temporal observations from other space-
and eround-based observations, would create a dataset of extensive coverage and variety^
These data could then be used as constraints on theoretical models quantifying the
ohvsics of the large scale global corona. One such analysis has been proposed by
Eiesecker and Gibsln in SOHO JOP 44 (for a full description see JOP044: Structure of the
Solar Minimum Corona, WWW), which would provide a quantitative description of the
global magnetic field - something that observations alone cannot establish. By combining
SeoryTnfdata we will gain a picture of the solar corona from the solar surface to the
interplanetary medium. [foP 04f-199608.txt, WWW]

This reference goes on to describe the magnitude of such a proposal. The main point
concerns the combination of information - theory and data, space- and ground-based
observations - to achieve better understanding. The fields of solar physics, solar wind
physics, magnetospheric physics and ionospheric physics have developed substantia y
using space-based observations. However, until recently, there has not been a concerted
effort to integrate these fields [Akasofu, 1996]. One recent effort is the ' Solar Information
Center" at Stanford University [Solar Information Center home page, Stanford
University WWW] which itself claims to be "under prototype development and only
exists in very rudimentary form." There is also the International Solar Energy Society
(ISES) which co-ordinates data collection from 10 Regional Warning Centres (RWCs)
throughout the world. Each RWC is funded by its host government for its own solar
warning purposes but the data are also sent to the U.S. RWC in Boulder, Colorado which
then collates them and issues world-wide warnings. Co-ordination of solar science data
measurements, at a certain level, is also already achieved through the Inter Agency
Consultative Group (IACG) [Johnson-Freese, 1992]. Also, " So1 ^ Posies data occupy a
sizeable portion of NSSDC’s archives" [Solar Physics at the NSSDC, WWW]. So, there
appears to be decent co-ordination of current solar data within the solar-terrestrial science
community. However, due to the wide range of global effects [section 4.5] there is the
need to make solar data more easily accessible beyond the traditional solar related fields
into the areas of climatology [section 4.5.1], sociology and medical research [section 4.5.2]
and technology fields [section 4.5.3]. The co-ordination of solar data should also include
those data from the past. It is obvious from the list of past and current missions [section
2.2.4] that careful organisation of the existing and incoming data is essential if we want to
exploit these data to extract as much information as possible.

Much solar data are remotely-sensed observations of electromagnetic radiation, and
therefore co-ordination opportunities are the classic ones faced by Earth remote-sensing

observers:

• it is more efficient to avoid similar observation and acquire data from different
temporal, spatial, and/or spectral areas,

• different temporal, spatial, and spectral observations can be combined to
produce much more information than the sum of the three individually,

• existing and/or historical data sources may prove to be complementary to new
data sources, and

• in-situ observations complement remotely sensed data.

Based on the above, the framework which has been set up through Earth observation
networks can serve as a model for solar observations.

onn • Ra: The Sun for Science and Humanity

8.4.2 Means of Achieving Co-ordination

h!iH P \ the ea ! ieSt su 8f stion and ma y be most important, is to continue with ground
based observations and ensure that these observations are accessible and, Indeed

Tu! a T l r S ° f Space ' based observations. These data are already being
al 19961 3 A d/ thr rf: heir ° nger hist0ry ' P rovide solid foundational data [HofLan ft
u l9 iA 6 u A y eff ° rtS t0 mamtain co-ordination and accessibility of ground-based data
should be encouraged. This will help support both science and applications ogives by
providing Ground based observations which are virtually continuous and provide the
low spatial resolution "big picture". p e rne

F - a f c °od suggeshon, borrowing from Earth Observation, we suggest the publishing

n ir a pr f Un t 0 r b w e 7 ah ° n Director y" similar to the "1995/96 Earth Obsfrvation Spacecraft
Directory Matra Marconi, 1995] which is a pocket sized directory updated semi¬
annually. (We have left out the word "spacecraft" for the Sun observations booklet
because we believe ground based observations should be included.) As markets increase

[ AkasoflTl 996 H tw^l 3 ^ h § 1^6 7] *** ^ Scientific dis cipHnes become more related

[Akasofu, 1996] it will be helpful to have an up-to-date directory of solar observers This

inexpensive and relatively simple suggestion could prove very helpful in organising solar

data sources, especially for those not from the traditional solar-terrestrial fields This

suggestion is most helpful for meeting those application objectives.

tuft SUgSeSti T t0 WO [ k t0 6Stablish an international Solar Data Centre" for both
solar science and solar applications as well as other fields which may be interested in

exploring solar data. From an international perspective we see the most efficient use of

° ar , data 1S to , hav ® the data available to as many users as possible. Our suggestion is

effort noTa dlta T* ^ be 3n international data co-old,nation

s? ,o be,ier “ studtes ** - -«

As an example from Earth Observation, a WWW Search for "Earth Observation Data

Observ a tion^ETwork^s°E SI ^!f Otf ^ ™ °t? hich ' as an example is Netherlands Earth
Observation NETworks Earth Observation Data Center, United Kingdom, which allows

WWwT H br ° WSe ° n keyW ° r f d and location [Earth Observation Datacentre home page
WWW However, a search for "Solar Observation Data" found no sites. It would be
helpfu 1 to create a browse-able international network for past and current solar data

university, 1 WWW]liaTstarte^tifdcx) ^ “

A fourth suggestion is to encourage researchers to investigate all possible data sources
mcuding those from the past, and have those data sources be relatively easily accessible
to the scientific communities (which, of course, would result from the realisation of the
previous suggestion). We recommend agencies consider using grants and fellowship to
.rnfate research which integrates various solar data, current ancffrom ^ev ous^s Ls

ICu 19961 ‘This ^ eXiS h ng - Slmilar “> lhe —1, diseased by

lAkasotu, 1996]. This research would help maintain our ability to absorb and
comprehend all of the existing and proposed solar-terrestrial data. We recommend this
o be an international programme where the amount of funding in research grants and
fellowships given m a participating country would be proportional to the amount of that
country s contribution to the programme. This would increase intematona co
ordination as well as provide efficiency by having a common administrative unit

Another example from Earth Observation is the North American Land Characterisation
Program which has organised “triplicates" of remotely sensed imagery for die^amearea.
one S from the early 1970s, one from the early 1980s, and one from the early 1990s [North
American Landscape Characterisation, WWW]. This type of "value added data package
allows the user to focus on the content and not the gathering of data. We believe that a
"Solar Observation Data Center" could facilitate some initial data processing to prod

enhanced data products.

A final suggestion for this section is to assume co-ordinated data access for future mission
there is some risk in such an assumption, we trust that co-ordinated
effort will always be most efficient. Assuming co-ordination in all future missions wi
more or less force co-ordination - because there will be no other choice. For solar-system
space science, international collaboration "has been outstanding' and is a given [ a e,
1996] Solar science and solar-terrestrial science would be wise to follow this example.
Working toward, and then assuming, international co-ordination will help achieve
observations from as many temporal, spatial, and spectral areas as possible.

8.4.3 Summary of Recommendations on Co-ordination

The near-term recommendation for data co-ordination can be summarised as :

• continue with ground based observations,

• publish a "Sun Observation Directory"(pocket-sized),

• develop an international data centre,

• provide support for research which co-ordinates scientific data, and

• assume data co-ordination in future planning.

8.5 The Near-Term Role of the Working Group for International
Solar Exploration & Application (WG ISEA)

The WG ISEA is a recommended framework to act as an international forum for the
planning, co-ordination, and implementation of an international effort in solar
exploration and applications. To do this, the WG ISEA is structured to incorporate
representation from both government and private sector space science and applications
interests as they pertain to the Ra Strategic Framework [section 3.2]. The changing globa
paradigm for space science and applications points to the advisability of combining
resources across both national boundaries and science vs. applications disciplines. The
Ra team believes that the WG ISEA represents the most efficient and expedient
organisational form to enable this merger for the benefit of international solar study
efforts Specific recommendations for action from the WG ISEA to its member agencies
should form the basis for an international collaborative effort in solar exploration and

applications.

The Ra report follows a phased approach in which each subsequent period builds on the
one before it. It is important, then, that the multi-lateral planmng, co-ordination, and
implementation effort begins immediately. While in the near-term Ra reconunends no new
flights (save for Cluster recovery) the need for the WG ISEA is immediate for a variety of

factors:

• Multi-lateral co-ordination of data sets from current spacecraft and projects such
as they pertain to Ra is needed (such as appropriate military satellite data sets as

Ra recommends)

i -Term Mission Scenarios require advance planning and budgetary
designations in space agency funding cycles. It is necessary for the WG ISEA to
su nut its findings and programmatic recommendations to agencies before the
budgetary cycles for the target years are "locked in."

The culmination of the ISTP, combined with NASA's Sun-Earth Connections
P f nmnS u make the P resent a un >q ue Period in space science for solar
wr rccrfu ph ! yS1CS , and a PP llcati °ns-a uniqueness of which Ra and the
G S ^?^! d T ^ ke advanta S e - In order to maximise its influence on this
penod, the WG ISEA should be formed and active before NASA's planned

WWW] H ° 6 meetmg m the summer of 1997 [SECAS Roadmap Planning

8.6 Increasing Awareness

Clearly the Sun is the most obvious celestial body. In any society, developed or
eveloping, people can idenHfy the Sun. Many people enjoy the peaceful and wonderful
experience of watching a sunrise or sunset. Yet the dynamics of the Sun are not very
pparent when one sits on a beach. Most people have the opinion that the Sun is a
relahvely stable fiery ball far away from the"Earth, while it is generally obvious that
arth receives heat and light from the Sun, solar physics beyond the photosphere (for
example, CMEs and the intermixed Sun and Earth electromagnetic fields) are not yet part
of what we could call "common knowledge," even in more developed societies. Perhaps

other an!i ^rh 18 '** dramatic effects of ,he Sun and their interactions w.th each

19961 hr t£ Sa , “ are ,, n °‘ we " understood mn ‘he space community (Worden,

1996J. In tins section we will discuss why it is essential to increase understanding of solar

understanding ““ SU »S esti °" s -create to

8.6.1

Need for Increased Awareness of Solar Physics

The Sun makes an excellent case for the complexity of nature and the nature of science-
different observations provide new clues to help our understanding and the phenomenon
is certainly more involved than what is apparent from what we see every day What
could be a more effective and interesting way to explain the range of the eTectromagnetk
spectrum as well as the complexity of solar activity than to display X-ray and ultraviolet

199?!! R f ^A Se t mOSt 9ny ° f the soIar * related WWW sites; Lang and Kenneth,
5, or Beatty and Chaikin, 1990, p.25]. Indeed this is why we, the Ra team chose to use'

hunums°the SOHoTJ' “ Trf °7 he , Sun ‘ aken by “ eye nera Possessed by

™ ' * SOHO E ^ treme Ultraviolet Imaging Telescope. As the solar-related fields

continue to grow with new observations and new theory, humanity too can grow by
sharing in the complex and amazing knowledge of our Sun. 8 Y

8.6.1.1

Need for Increasing General Public Awareness

Most space exploration and science is publicly funded. The public tends not to want to
pay for something it knows nothing about. Space programmes must now realise tha°
good science is not enough to keep a programme funded [Randolph, 1996], Currently
budget constraints require space agencies to pick and choose programmes carefully. Any
programme needs to justify its spending, not just to scientists in the field but also to
politicians and to the general public. This implies a need for communicating the
importance and relevance of space programmes in common language. 8

Not only do space programmes need to justify their budgets but also they should spark
the interest of the public as well as share the importance of their findings. Science,

technology and space exploration affect all humanity : they help set our course for the
future and reflect^the general human endeavour to explore. So whether it is to work
toeether to plan the future, to share in the excitement of new findings, or to understand
how^ax^m^ney is being spent, as much of society as possible should be aware of
advances in space exploration, science, and technology.

8.6.1.2 Opportunities in Education

We consider educational opportunities a strong component for increasing public
awareness and involvement with space exploration. Space is an inspirational tool for
science and mathematics education [ESA SP-384,1995], A1 so, asMn se hon 4.1
we know there is a social component to our explorations, and from chapter 3 there a
related policy issues. It is the many facets of space exploration which make it an excelle
StU* students of all ages. This need has beenwell ^

incorporated into the outreach programme accompanying the IMAGE spacecraft,
approved in 1996, which states :

The IMAGE Mission Team will be involved in a program of Public Outreach
Education, Teaching, and Reaching Youth : a program we call POETRY. IMAGE
will Droduce spectacular images representing the plasma environment of the
Eartlf These images will not only allow the IMAGE investigators to understand
The physics of the magnetosphere, but will entice the public, students and
testers into learning more about the fascinating and complex processes that
surround the Earth. [The IMAGE Mission : Imager for Magnetopause-to-
Aurora Global Exploration, WWW]

8.6.1.4 Opportunities for Sun - Earth Interaction Awareness

Everyone connected to space exploration should, at the very least be aware of the effects
of solar activity on spacecraft [section 5.2]. One-half to three-fourths of anomalies m
satellite behaviour are correlated with space environmental disturbances [Worden, 1996].

It is impossible to say whether the space environment has caused these anomalies
because P at present, there has not been substantial research m the literature which
established a^orrelation between solar activity and spacecraft malfunction. There is litt e
connection between the government solar warning services and their users. So what if
my satellite or power line or pipeline is hit by a geomagnetic storm? Does anyone know
what I should do about it? At the present time, not really. There are some attempts to
develop operational models for various users to instruct them on what to do when certain
dangerous phenomena occur but they are far from complete. Satellites aje certain y not
designed with these phenomena in mind. There exists a weak link that must
improved between warning services and users.

8.6.2 Means of Increasing Awareness

Reflecting some of the different opportunities described in the previous section, we now
offer some recommendations to increase awareness.

(See [ESA SP-384,1995] for a general approach to communicate space agency activity to
society.)

8.6.2.1 Awareness for the General Public

Scientists should be either encouraged or required to make at least the essence of their
findings available and accessible to the general public. A requirement would probably
be mef with reluctance or even strong resistance. The best direction to take is probably a
"strong encouragement". As an example, funding agencies could suggest that each
technical proposal and research report derived from work sponsored by that agency be

,Z Pa 7 ty 3 sm 't’ l,fu:d document (of the size of one to two pages) written in

resTaT T^ UagC ' " hich relays ,he ‘"‘cresting and/or fundamental elements of the
research. This ts sinu ar to a press release, but at a micro scale for any reZch acHviW
and research report, ft should be possible to describe even complex cSt 1!
common language. Having scientists do so will give space agencies baseC ^orma Ho

Ho dd mUS6UmS ' the media ' a " d <°publifh on the WWW “d

star Id 6SS "X f ° r ata " dardisa *“>" on certain scientific terms referring*. diffeem

solar and space weather phenomena and regions of interacHon. 8

agencies t addiHont lmpo ‘ tance of P ubllc outreach, funding agencies (national space

in the Un tedifaTes?should n He ng ° r « a " ,sa,ions such as * ha National Science Foundation
scientist^ (and „,b H d pr0m0,10n opportunities, grant requirements, etc., into a
nt, " 1 * <and other space activates participants') public outreach performance and
p ans. The key word here is performance, versus mere effort. It should be a responsibility
of scientists in space acHvihes to pro-acHvely address public outreach and education and
this means improving their public communications We are not proposing hat' this

™ ttr “““ ? fUnding aVallabUi, y a " dy "‘ P—n opp~
„bL, be consldera ‘ lons taken in account in the decision making process The

o°u'he C aIin S aC ‘ ,Vity Partk ' Pan,S """ ° f tha “"Pounce 0 o e fpub T hc e

8.6.2.2 Awareness through Educational Programmes

training packages for instructors) offer an effective way to communicate the ever

/access tenter - NASA Observatorium, WWW1 thp "Qnlar » i 6 .

page (Solar Connections, WWW), and "SpaceLinlC IKASA Soarf t”' a ed " cal,onal

educational materials programmes, should T in ^"1

experiments 3 ^ reSe3rChinStiluHonsand beused toe " ab| ostudents todoXhownsoto

We applaud the plans like that described for the IMAGINF a

andTaMer^" 08 " 31 ^ 65 t0 ^ induded in a11 s P ace science mission plans ~ near^ mid- 6

8.6.23 Awareness of Sun - Earth Interactions

It seems that the best opportunity to educate the space community and power comoanies
is through a correlation study which aims to relate satellite malfunri^n T ,

activity. If, indeed, these are found to be related it would become clear to thos^' T & a

fhaf „ is in fheir in,arcs, ,o learn more abou, the effec, ouZ atvi^

vehicles and, then, work to develop both operational procedure and future engineering
developments which could reduce or avoid damage.

SSSrSESSSsSSsSS

With this' it would be wise to invite outside consultants. The space agencies can play
m e or Sthe study and disseminating the results. The space /government
agencie°aruTmilhary organisations can provide data (primarily on solar activity) and
expertise The satellite producers, communications industry, and power co p
would also contribute data would be

comprised 0 of 8 a tTaTwhich, collectively, had knowledge of solar-terrestrial physics
possible effects on satellites and power sources, and sufficient statistical h a ckground to
conduct a time-series correlation study [Detman and Vassiliadis, m It wou

informative to look at the anomalies with respect to the damage (if there was any) in
terms of cost. Any significant relationship should prove interesting and educatio .

8.6.3 Summary of Recommendations on Awareness

The near-term recommendation for increasing awareness can be summarised as

. requesting and organising “common language" summaries for science reports,
and makfng public outreach performance and plans an evaluation tool in
funding and promotion determinations/

. space agencies, and possibly commercial educational resources, working with
educators through the WWW, video productions, and workshops,

• a correlation study on satellite anomalies and solar activity.

8.7 Actively Incorporate Existing Technology Initiatives

We believe that in the near-term, and through the far-term technology development
should follow the "faster, cheaper, better" approach because doing things slower,
expensive, and worse" would be wrong. (But seriously...)

A growing international trend that has emerged during the recent years, is a push for
"faster cheaper, better" (or some other permutation of that order) space programme -
both civilian and military. The forces driving this change are mainly the pressures of
declining budgets in the post-Cold War environment, the emphasis on reduced
programme rifks, the emergence of advanced lightweight technologies, and the
development of low-cost, small launch vehicles.

We believe that the Ra programme, from the start, should incorporate this technological
philosophy Smaller satellites mean simpler design, smaller launchers, smaller
management organisations, shorter development time, and hence cheaper and ultimately
more missions With faster missions, there is greater opportunity for the incorporation of
state-of-the-art technologies, and there can be an improvement in technology based on
the flight results. In addition, if the mission development time is short (i.e. not a decade
like previous spacecraft development times), participants can be involved in all phases of
the mission, and personnel morale can be maintained.

The faster cheaper, better paradigm has proven to be a successful one in securing
government funding, and is espoused by NASA, ESA, and elements of the U.S militarv
space programme. Recent and current examples such as the Clementine and DC-X
demonstrate the utility of this approach in achieving results, and NASA is
basing its New Millennium missions on this approach.

°/h““^ here are ar « uments that " s ™"" ™»y not necessarily be synonymous with
cheap . There are two categories of "small" : y

1. simple spacecraft with fewer functions based on a standardised bus and off-the-
shelf components with minimum performance,

2. the miniaturisation of conventional components using new technology with

high performance in mind. hy

^Tjrrr ( S m ° re C ° Stly ' u Ut tHere iS increasin S interaction between the military
sector that tends to favour approach second within the civilian sector. Hence, technology

ransfer is something that must be encouraged to continue, so that the civilian sector can
take advantage of more performance oriented technologies. In summary, the advantages

n _ r ?. faStef/ cheaper ' better approach are numerous, and given that this current
p adigm has been and is successful, the Ra programme must foster, encourage and
incorporate this philosophy from the very beginning. '

8.8 Conclusions

We believe the recommendations are realistic and play an important role in realisms
important science and applications objectives. They also provide a foundation for thf
programmes described in the Mid- and Far-Term components of the Ra Strategic

Chapter 9

n r>

1

^ Introduction A

f

^Our View of the Surii

f

(

V^eeds & Objectives jj

1

<==> Cl c*' f I

^Strategic Framework Jf »

j-

9

nUT:

C Technology A

frSV

^ Near-Term A

ifKT

Mid-Term j

ifSKi

Far-Term j J

V^t arke ti ng^^u nd i n g J

..mill'

. I

| Political & Economic"

*

V Environment J

/ijm

t "Conclusions* - f

Mid-Term Programme

The part of the Ra Strategic Framework comprises a number of suggestions

including a solar monitoring and early warning system, a pure applications mission as
well as a dedicated science mission to study the Sun from 0.2 AU.

9.1 The SAUNA Mission

The SAUNA (Solar Adjacency Using a New Approach ) Mission is a new system to perform
solar scence m a low solar orbit over a time span of several years. Unhke e g the HRE

«™a sna A ft d °n n0t a ', ,empt a Slngle «>™ a Instead the

AUNA spacecraft will go as close as requirements for a multi-year lifetime allow.

SDcKrecraft^tia morUtn" tK* ere ^ also . serves as a demonstrator for the constellation of
(Far-Term) the Solar envlronment in this region, as described in chapter 10

We have endeavoured to make this mission politically acceptable by designing to a US$

00 million Life Cycle Cost and by applying no controversial technologies like
Radioisotope Thermoelectric Generators (RTGs). ° Ke

9.1.1 Design Procedure

The SAUNA design has evolved through the following process:

• Mission requirements definition

Mission feasibility: Trajectory and propulsion studies

• Establishment of preliminary budgets

• Spacecraft configuration trade-offs

• Subsystem design and sizing

The work was fast-paced, of a parallel and interactive nature.

9.1.2 System Architecture

This section describes the system architecture of the SAUNA mission.

9.1.2.1 Mission Objectives

The SAUNA mission will perform in situ scientific measurements from a near Sun orbit
The scientific measurements will focus primarily on studying the solar wind and the
Sun's corona and secondarily on studying interplanetary dust. Finally the mission will
prove the survivability of a spacecraft in a near Sun orbit. This las. objective is regarded

as primary.

The study of the solar wind will be focused on analysing the solar plasma and measuring
the Sun's magnetic field. For studying the Sun's corona the Sun s photosphere and
chromosphere^will be studied in extreme ultraviolet, energetic particles will be detected,
and solar eruptions and coronal structures will be imaged. Interplanetary dust will be
studied during the transfer from Earth orbit to Sun orbit. These objectives will be further

discussed in 9.1.4.

Survivability in a near Sun orbit will be considered proven if the spacecraft provides
prelection Lainst the solar environment to the extent that the payload rema ns
operational during the required mission lifetime. Protection against the solar
environment in this § context includes maintaining a proper orbit and attitude, a hermal
environment in which the spacecraft systems can operate, a communications link to Earth
that enables a specified data volume and flow, etc. If for instance an instrument s
destroyed by an impacting meteorite (a generic threat in space and an unavoidable risk
for exposed instruments), this would not mean that the survivability objective is not met.

p— Study Solar Wm3

L_3Pdy the SuiVsCocna.

SAJNA J

-c

Analyse Plasma

Jjtidy

nterplanetaiy Dust

14 agnelc Field
VI easu aments

—Exfreme Uhrauolet Imaging
^neigetic Particle DetecHon
—Doonogrtphy

0 emo n si al e su vebi I ty i n
Vow Sun Orbl

Fig. 9.1

Overview of SAUNA Mission Objectives

9.1.2.2 Mission Constraints

While being in Sun orbit, the SAUNA spacecraft will be operational for a minimum of
three years Furthermore the study of interplanetary dust, which is a secondary mission
objective, may not interfere with any of the primary mission objectives. Finally the total
mission life cycle cost shall not exceed US$ 200 million. For the total mission life cycle
cost scientific analysis of the data is not taken into account.

9.1.2.3 Functional Analysis

Figure 9.2 shows the top level functions for the SAUNA mission as depicted in a
Functional Flow Block Diagram.

210 • Ra: The Sun for Science and Humanity

JPerform Pr*-I»unch I
10ptration* , f

I Perform Laurrch
J Cp*rations ? I

I Partorm Sun
lScunci Wisjio

I Perform End-0< Lif*l
Operations f. r

I Partorm Interplanetary I
1 Science Mieeiori . P

Fig. 9.2: Functional Flow Block Diagram for the SAUNA Mission

In this figure 'Perform Sun Science Mission' refers to the study of the solar wind and the
d!T“ Terf ° rm Sci — Mission' refers to the study of

For the SAUNA mission the following top-level functional requirements were identified:

• The SAUNA spacecraft shall be accelerated beyond Earth escape velocity

• The SAUNA spacecraft shall be transferred to a near Sun orbit

The SAUNA spacecraft shall study interplanetary dust during the transfer
to a near Sun orbit

• The SAUNA spacecraft shall establish an orbit around the Sun

The SAUNA spacecraft shall study the solar wind during its stay in Sun

• orb G it SAUNA SpaC6Craft sha11 stud y the Sun's corona during its stay in Sun

The SAUNA spacecraft shall provide an operational environment during
the mission lifetime 6

9.1.3 Mission Design and Spacecraft Configuration

their outers 118 ^ dlSCUSS mission desi S n ' spacecraft configuration studies and

9.1.3.1 Mission Design

The focus of the mission design work was to define a low solar orbit and a transfer

the^ r aWe '° '“I" * he miSSi ° n 0b ' ecHves ' As a " ‘"put these problems

the following factors were considered: '

Target Orbit:

• The mean distance from the Sun would have to be small enough to be

T 3 standpoint, yet large enough to sustain an

extended life (thermal and radiation environment)

• The orbit would have to be attainable for a small spacecraft using presently
or near-term available propulsion technology

The cost of getting there and staying there must be balanced against the
cost constraints of the mission.

Mid-Term Program mp • 911

• The transit time should not exceed the time spent in final orbit

. The selected trajectory should not put extreme constraints on the
deliverable dry mass

• The propulsion technology and performance required should be available
in the near future.

• The injection conditions required should not lead to extremely high launch
costs or assume the use of currently unavailable launch systems.

Initial trajectory studies quickly led to the dismissal of chemical propulsion for the
transfer orbit due to the low I sp , leading to a very small mass fraction. The circulansation
of an orbit near the Sun requires a significant Av which proved to be a very diffic
condition to meet. Solar sailing was excluded as an option due to the immature status of

this technology.

We selected ion propulsion, as this technology best covered our mission needs. The high
I and low thrust leads to a high mass fraction and (relatively) long trip times. Using an
ion engine with 0.2 N thrust and an I sp of 4700 s [section 9.1.5], we assumed an initial wet
mass of our spacecraf. of 300 kg and arrived at a feasible trajectory to a 0.2 AU circular
orbit [figure 9.3], by using the SKYNAV optimisation software [Appendix CJ.

Low Thrust Trajectory from 1.0 AU to 0.2 AU Circular Orbit: In Bold Lines
is the Thrusting Phase of the Trajectory.

The total transfer duration is 507 days of which the first 90 days are spent coasting on a
Venus transfer trajectory (this requires a launch energy of at least C 3 - 15 km /s u¬
rest of the transfer is a continuously thrusting manoeuvre. The transfer time, although
long does not exceed the minimum survival time expected of the spacecraft in orbit.
Furthermore, we have not included the effects of a Venus gravity assist in the above

result.

We do not pretend to have an optimal solution to the selection criteria described above,
but this trajectory does meet all the requirements.

9.1.3.2 Spacecraft Configuration

* T*-spacecraft muat be provided with a heat shield to protect against the
thermal and radiation environment close to the Sun.

' I h J\r‘ a ; aI L a ? S mUSt be Sized 10 P rovide the squired power throughout
the transfer trajectory and when in orbit, while the solar energy flux dfnsitv
increases by a factor 25 from 1 AU to 0.2 AU. ^ aensit y

When near the Sun, the solar array must be protected from the heat to

theln^ an ,^ XCessive rate of degradation and to ensure that the bonding of
the solar cells is not compromised. s or

* The ion engine must be located on the forward part of the spacecraft to

provide the retrograde (braking) Av. spacecratt to

* transmission" ante " na mUS ‘ * P ° inted tOWards ,he Earth for hl 8 h data

or 0 Do!ntaro P f a r y hr d 7 quireme " ,s were ind “d«l i" the considerations above (e g FOV
ins^ent? could bT^nd

requirements 3 ” ^ 0d ’“* r “ ilh « <**■*■*>« row”* fhe p^.oad

Choice of Stabilisation

T . h f® 1 ® ^ pes of stabilisation principles were evaluated for SAUNA- Thr«a 0

spin, table 9.1 was used for evaluation: ' 3 stabilisation versus

Table 9.1 Evaluating 3-Axis vs. Spin Stabilisation.

The choice fell on 3-axis stabilisation mainly due to a consensus that tho a a

---

Selected Configuration

central enclosure accommodates the

JSX "o^etn: 4"e ^Tnglne is Lusted on the forward (velocity vector,
side of the spacecraft.

SUN

Heat shield

Velocity vector

Magnetometers
on booms

Fig. 9.4 SAUNA Spacecraft - Selected Configuration

The functional concept of this

opttma 1 pomtmg o^ th filing over so that the array is perpendicular to the

engine. Ttos can be ach J »’ hes the Sun, the solar energy flux increases

incoming radiation. As the spacecmn pp^ ^ ^ ^ ^ effects; (1) The

inversely proportion^ electric power generated by the solar array increases

rrr se ts u By g^y m

Wktofero degree roll angle, the spacecraft simultaneously increases its thermal
protection from the heat shield and ensures a bounded power output from the solar
array. Figure 9.5 illustrates this concept:

Fig. 9.5 Spacecraft Orientation (a) Far From the Sun, (b) Close to the Sun

changes. This principle is illustrated in ^ enablmS eg ' ° rblt inclination

SUN

Fig. 9.6 The Thrust Axis Can Be Pitched To Provide Out-of-Plane Av.

Launcher Considerations

transfer orb,, on

;“h“a-r^ WU1 haW ‘° ^ deP '° yab,e Mde;S,0 “ I-

9.1.4 Science and Payload

This section describes the science background and instrumentaHon payload of SAUNA.

9.1.4.1 SAUNA Science Background

measurements, from" ^erat^^

pro^r.o“ m 8 ;zxii n wii ;

resolution. The improvement in resolution will h» a JL P aIu te , m P oral and spectral
probity ,0 the Sun and i m “t^^

missions. ,n si.u measurement wiU provide, informaHorton the."r^solar
wind in a region not Piously sUtdred overmuch aceof theSun(especiaUy

and J l

”om sXutJr"combin^i wfth those from spacecraft in the vicinity of the Earth to
obtain stereoscopic and contextual information.

This data set will be extremely useful for ^trTsuch a°s

nudKf—^lH,s S t£ moSng and prediction of solar processes which
affect the Earth.

914 2 SAUNA Instrument Package

The payload of a Sun — »,

least, a white light coronograp . . that none G f the existing instruments

AWA Cia"“ e^are separate and hence

jk em'ssx

longer life time (it will cause increase of weight).

,. . £ ll« c 1ir t if would be very desirsble to hcive 3 pointing

around the heat shield.

The estimated characteristics of the considered instrument are based on the enumerated
instruments as well as peculiarities of the SAUNA orbit.

• Visible light coronograph: 10° FOV

• Ultraviolet:

1) 4° FOV - to monitor the solar limb

2) < 0 5° FOV - to monitor specific areas on the solar surface with high
resolution benefiting from close proximity to the Sun.

Such an instrument would consume 5-10 W of power, weigh 6-7 kg, produce an average
data rate of 5-10 kbps and cost roughly US$ 10 million.

relatively high mass and can not fit„ , " , ° n * e tQ current NA | A and ESA study

requJste'could meit a SAlSA requirements [NASA Research Announcement, 19951

We conclude that the fo^itT ^un^orbUm^phTse^plus'^dust^detecUir^or

measiirements during 3 the fransfer Rectory phase. The total weight and power of the
instruments will be of the order 15 kg / 40 W.

9.1.5 The Propulsion System

I n J h t t f0l L°r; 8 W d ? Cr ‘ be ,he P r °P ulsion s y stem ’ wi* emphasis on the main engine
and the attitude control engines. &

9.1.5.1 Main Engine

As mentioned in the mission design section, we selected a high thrust ion propulsion
engine. This engine has a real-life counterpart: The UK-25E thruster [Latham et al 19951
w ich exists as an engineering model but has yet to be flight qualified. Its high specific
impulse provides a good mass ratio and saves propellant.

Table 9.2 UK-25E thruster [Latham et al, 1995]

Nominal

Thrust

206 mN

Electrical

Efficiency

Propellant

Operating

temperature

Specific Impulse

6300 W

The thruster has a design lifetime of 10,000 hours (converting by coincidence to the exact
time of thrusting for our selected transfer trajectory). It has a mass of 20.6 ke and will
cost approximately US$ 200,000 [Martin, 1996]. 5

Problems

• The engine has to be qualified for at least 10,000 hours continuous thrust

• The frontal engine mounting leads to plasma backflow during thrusting
phases (potentially damaging to solar array and instruments)

Attitude Control Thrusters

Main Engine

Fig. 9.7 Propulsion System Layout

9.1.5.2 Attitude Control Engines

We T? 6 u ma “ ,hruslers (required for ^ stabilisation) [Larson &
Wertz, 1992], They will also be .on engines which use the same main tank for their fuel.

his saves mass as compared to separate systems for attitude control and main
propulsion [figure 9.7].

9.1.6 Power Systems

Table 9.3 Solar Arrays [Larson and Wertz, 1992]

Source

Efficiency

Required power

Array

size

Problems

Counter¬

measures

GaAs Solar
arrays

19%

7 kW (incl. 10%
margin and all
sub-systems)

14 m 2

Temperature at the
Sun - bonding of the
arrays

Tilt arrays

Degradation

Balance
losses and
increased
solar flux

The solar array consists of a single fixed tilted solar array
the Sun by rotating the spacecraft. We did not make a
regulation system.

which can be oriented towards
choice of batteries and power

Suitable high temperature bonding agents will be used in the development of the solar
array.

9.1.7 Spacecraft Structure

The precise geometry, size, materials and mass of the SAUNA spacecraft structure have
not been defined. The primary structure might be made mostly of Titanium alloy,
ceramics would be used for those elements most exposed to heating (e.g. heat shield

support structure).

Critical problems include the moving parts such as deployment mechanisms for the solar
array and the antenna. Very high reliability will be required, and this increases
development costs substantially. Continuously moving parts and rotating joints such as
the antenna pointing mechanism will require special lubrication and tribological
measures to avoid cold welding and potential malfunctions.

The spacecraft structure as a whole must withstand the loads and vibrations induced by
the launch vehicle. Corrosion induced by the ion engine plasma may have to be
counteracted by special measures.

9.1.8 Thermal Control System

In this section the thermal control of the SAUNA spacecraft is discussed. Special
attention is given to the heat shield and the thermal control of the instruments and ion

thruster.

9.1.8.1 Thermal Environment and Requirements

When the SAUNA spacecraft is in its target orbit, the distance between the Sun and the
spacecraft will be 0.2 AU. The heat flux of the Sun is therefore (1/0.2) =25 times higher
than at Earth. Moreover, due to the solar wind (the flux of particles ejected by the Sun)
the frontal surface, which is always pointing at the Sun, is continuously subjected to

particles.

We consider only the heat input by the Sun and by the spacecraft systems (a really close
encounter with Mercury is not likely). The thermal production of the spacecraft consists
of two main contributors:

• The instruments, communications system and the on-board computer
which dissipate in total approximately 100 W.

• The ion thruster system of 6.3 kW, which dissipates (Worst Case) 1.8 kW of
heat (with special requirements for propellant tanks and batteries).

The temperature requirements for the spacecraft components are listed in table 9.4.

Table 9.4 Thermal Requirements for the Components

Component

Temperature Range f C)

Electronics and Science Instruments

-20. .60

Batteries

5...20 |

Ion Propulsion

300...400

Xenon Propellant

>20 at >125 bar

Structures

-45...65 1

Comparing the temperature requirements and the dissipation, it is clear that the thermal
control system must be divided into two systems., one which takes care of the low
dissipation sensitive instruments and electronics and one which takes care of the 1.8 kW
heat dissipated by the ion thruster system.

9,1.8.2 Thermal configuration

The solar radiation (at 0.2 AU) is a flux of about 34 kWm 2 [section 6.7]; using a heat shield

spacecr S aft h 1rom P thl 0f I*' 5 ° D ^ s P acecraft Moreover, it can protect the

p cecraft from the solar wind. Various materials can be used for the heat shield

ITowever, Carbon-Carbon is up to now the most promising candidate [section 6 7] It is a

we 1 known material (for temperatures below 2000 K) and it has a solar absorption and

emission coefficient which is, comparing with other candidate materials, insensitive to

impact of UV-radiation and solar wind. It is expected that the outgassing of a carbon-
carbon heat shield is not a problem because: 6 8 <-arDon

The solar wind will interact with the outgassing atoms and therefore cleans
the surroundings of the spacecraft. This reduces the danger of
accumulating gases surrounding the spacecraft.

• The outgassing rate reduces with an order of magnitude for every 100 K

?nnnt G °^ aSSm Z rate is about 2 m gs' [Millard, 1992] for a temperature of
2000 K. Thus it is expected that for a temperature of 600 K the outgassing
rate is worst case 2x10 10 mgs 1 . Taking in consideration that the total time
near the Sun for SAUNA (3 year) is about 2000 times longer than for the
Solar Probe [Randolph, 1996], at first hand, it is expected that the
outgassing phenomena will not have an influence on the plasma
measurements. K

^ k X from the Sun further ' a standard multilayer insulation (MLI) (e eff

- 0.015) is located between the heat shield and the instruments. The dissipated heat of
the ion thruster is transported to radiators which radiate the heat into deep space By
usmg a two-phase heat transport system, the temperature difference between the ion
thruster and the radiator is kept small. Therefore, the radiators can work at high

(c3“tod cX U toop? e needed SUrfaCe and maSS “ mpared 10 ° ther melhods

9.1.8.3 Heat Balance

In this section, a simple heat balance (steady state), for an orbiting spacecraft, is used to
determine the properties of the thermal control system. It is assumed tha .

. The heat shield is always pointed at the Sun providing shade for the whole
spacecraft.

• The thermal controls of the ion propulsion system and the instruments are
separate and independent.

The heat balance equation
temperature of the side of

feqn. 9.1] does not directly include solar radiation but uses the
the heat shield radiating to the MLI as a boundary condition.

Instruments spacecraft systems, and heaters are modelled in a box with one heat output
value, one teniperature on the outside of the box and one emittance value. The equation
is analogous for the ion engine (in the box) and its radiator.

Fig. 9.8

Thermal Model of Heat Shield, MU, and Instruments

Awzj^mzj

(9.1)

where:

A [m 2 ]

e

(= 5.669X10" W/(m 2 K 4 )
Q[W]

TIKI

hs

rad

General

a MLI =a . 5m 2

£mli

Ths

= 0.015
= 600 K

surface area (for MLIs: surface directly facing the heat shield)
emittance (for MLIs; overall effective emittance)

Stefan-Boltzmann constant
power expressed as heat flux
temperature
heat shield

radiation to deep space
Instruments

Arad

2

= 2 m

^rad

= 0.15

Ion thruster
Arad

^rad

= 1 m
= 0.8

Figure 9.9 shows the temperature of the instruments for various dissipation powers.

220 • Ra: The Sun for Science and Humanity

0 * 0 8 0 8
□uf*nc« »un[AUJ

Fig. 9.9

The instrument temperature as a function of the distance from
various dissipation levels.

the Sun for

It can be concluded that:

The heat shield temperature increases from 250 K, for 1 AU, up to 570 K, for
0.2 AU.

• The temperature of the instruments only varies less than 30 K for a distance
rom 1 AU to 0.2 AU. The thermal control can be passive by changing the
emissivity of the surface in the design phase. However, when the
instruments are switched off a heater must heat the critical instruments to
prevent too low temperatures. The heat loading structure can transport the
amounts of heat of the instruments to the surface of the spacecraft.

Figure 9.10 shows:

The temperature of the radiator, assuming that it is perpendicularly
oriented with respect to the solar radiation, is not sensitive to the distance

from the Sun for a distance > 0.2 AU.

• For the assumed radiator surface of 1 m 2 and emissivity of 0.8, the
temperature difference between radiator and ion thruster can be up to
about 100 K, which can be obtained using a two-phase heat transport

Fie. 9.10 The Ion Thruster Radiator Temperature as a Function of the Distance from the
° Sun for Various Dissipation Levels.

The batteries must be enclosed in a thermally controlled environment In the SAUNA
spacecraft, they get their own radiator area and thermostat-controlled heaters. The
thermal dissipation from the batteries varies with temperature, charge, and charge rate
and can be difficult to quantify. Therefore, the thermal control of the batteries needs
special attention in a more detailed thermal design.

The propellant tanks also required a tighter temperature envelope than the electronics
The propellant tanks contain xenon in supercritical state. Xenon must be stored a
>125 bar and >293 K. In the SAUNA spacecraft, the propellant tanks are m a thermally
insulated environment, with their own thermal control.

Table 9.5 gives a overview of the thermal system components and the mass and power of
the system.

Table 9.5 Overview Mass and Power of SAUNA Thermal Control Subsystem.

f Component Thermal control

Mass [kg]

Power [W]

Heat shield

8

-

MLI

3

-

Heaters

2

20

Conductive structure (part of main structure)

5

-

Two-Phase evaporator (connected to the ion thruster)

10

-

Radiator

5

-

Control electronics

2

10

Total

35

30

9.1.9 Attitude and Orbit Control System (AOCS)

The AOCS is composed of the following elements:

• Attitude determination performed by using the measurement from star
sensors and an Inertial Measurement Unit (IMU).

• Attitude control is provided by reaction wheels periodically desaturated by
ion propulsion thrusters.

999 • Ra: The Sun for Science and Humanity

he operation is as follows. The measurement from the star sensors is used durimr
nominal mode pointing towards the Sun to calculate the orientation of the spacecraff

used "J eaSUrement 1S aIso used to Periodically calibrate the gyrometer set drift that is
used during manoeuvres mode because the star sensor can not be used. Once the
orientahon is calculated by the on-board computer, the computer compares this attitude
vi h the assigned one and produces an order for the actuator to correct it The actuators
a?H “ WhedS Whkh WU1 reqUire 3 desat ^°n m °de in which the thru Trl Zl

The components of this subsystem can be the following:

Star sensor. The Star Tracker Stellar Compass (STSC) can be used. The
demonstrated accuracy is 150 prad in pitch and yaw and 450 prad in roll.
The weight is 290g, the FOV (Field Of View) is 28.9° x 43.4° and patterns of
stars as dim as mv 4.5 are measured and matched against an on-board star
catalogue. Star matches are achieved in 4ic steradians of the stellar sky
Two systems of this type will be implemented to achieve a redundant
systenu They will point towards a direction with 30° with respect to the
zenith direction (to avoid antenna, solar array, and radiator interferences)
lhe power usage is about 12 W [LLNL 1996].

IMU. It is composed of four HRG gyrometers in a tetrahedral configuration
to have redundancy. The drift is 0.006 °/h and the power usave is
estimated to be 24.6 W [Randolph 1995]. usage is

Reaction Wheels. They can be provided by Ithaco in a four wheels
redundant configuration. The momentum storage is 50 Nms (3850 rpm)
he maximum reaction torque is 0.3 Nm, the minimum lifetime is 8 years
the mass is 14.1 kg, the power at steady state is 35 W (3850 rpm) and the

power at peaks is 200 W (3850 rpm) [Ithaco 1996]. ^

Thrusters. Six thrusters are located around the spacecraft to desaturate the
wheels. More data is provided in the Propulsion section [section 6.4],

9.1.10 Communications

The SAUNA communications subsystem is divided into two different

parts:

• housekeeping communications

• science communications

9.1.10.1 Housekeeping communications

Housekeeping communications are carried out during all phases of the mission (cruisine
and orbiting). A set of 4 low gain antennas (LGA) in S-band is used in order to reduce thf
pointing requirements. The antennas are placed in different parts of the spacecraft in
order to allow communications regardless of the spacecraft attitude. P

This set of antennas could also used as a backup for the transmission of science data
However, only limited science data could be sent due to the extremely low Tata rate
achievable through the LGAs, 1.5 Kbps. y te

Appendix B contains the link budget analysis corresponding to the downlink of the
GAs, where the mass and power budgets are shown. The result is about 160 W (RF)
ivi 6 . ^ 4 power am P^hers, which result in an input power of 640 W (DC) This
amount of power is not a problem during the orbiting phasecdue to the large solar arrays

K/T i H -T<3 T-m Pr

The analysis has been carried out for the worst case in terms of distance and noise
temperature. Therefore, during the cruising phase the housekeeping communications
link operates reliably in spite of the reduced availability of power.

9.1.10.2 Science communications

Except for the cruise-mode dust detection experiment in transfer orbit (which needs only
a low data rate) substantial scientific operations are carried out only during the orbiting
piaTe In Ms phase the spacecraft is"3-axis stabilised, with the heat shield pointing
towards the Sun. A high-gain antenna (HGA) »» a diameter of 2 m operatmg m X-
band is placed in the umbra, pointing roughly towards the Earth.

To cope with the varying relative orientation of the Earth and the spacecraft along the
orbit the HGA needs a pointing mechanism. The need to use moving parts (bearings and
lubricants) eventually subject to an extreme and harsh environment complicates the
design of S the communication subsystem and reduces its reliability. Section 6.9 presents
some alternatives applicable to the SAUNA mission scenario.

The baseline configuration considers a single-axis pointing mechanism for the HGA^ The
motion along this axis (pitch) is limited by the spacecraft structure and the limits of the
umbra. The baseline considers a motion of ±90° around the zenith P° int 1S P r ° v '
coverage of around 50% of the orbital period, or approximately 15 days, the part o
orbit nearer to the Earth. The period of exclusion due to conjunction (15 , for safe

communications) corresponds to about 3 hours, within the coverage ^ol^Z sdencl
period, the radio system could operate in continuous wave mode as a plasma science

experiment).

If the SAUNA spacecraft goes out of the ecliptic plane (e.g. to a solar equatorial orbittorto
a higher solar inclination) a motion in the yaw axis must also be considered (2-ax s
pointing). For a solar equatorial orbit this motion is ±7°. The baseline is however not to

go out of the ecliptic.

The science data communications are carried in the X band (8.4 GHz). Appendix B shows
the link budget analysis corresponding to the downlink (worst case) of science
communications. The result of the analysis is 40 W (RF), resulting in an input power o

110 W (DC).

The link can operate with an effective maximum data rate of 16 kbps. The data are coded
using (255,223) Reed-Solomon block coding and rate 1/2 Viterbi (convolutional) coding.
Using this approach a bit error rate of 10' 6 is expected.

9.1.11 Command and Data Handling

The on-board Command and Data Handling (CDH) system consists of four main
elements: The main computer, the flight software (which resides on the mam computer),
mass storage, and the central data bus.

9.1.11.1 Main Computer

The Main Computer is the brain of the spacecraft. In terms of hardware it needs a very
fast processor for parallel and real-time operations at a high frequency:

• Execution of the GNC software

• Execution of the Vehicle Management software

• Processing of telecommands and packaging of telemetry

224 • Ra: The Sun for Science and Humanity

In chapter 6.10.3, autonomy functions for SAUNA-type missions are discussed.

9.1.11.2 Data Compression

The constraints imposed on antenna sizes and on mass and power budgets which are
especially important in the SAUNA programme due to cost constraints 'and to the
mission scenario, limit the data rate available for science data communications

capacity 113 ^ ** ^ iS ^ " eed t(> reduCe the on - b ° a rd storage

These reasons point to decentralised data compression. Compression ratios of up to 40T
can be achieved without significant degradation of the data quality. In SAUNA we
propose t e use of techniques providing an average compression ratio of 32 1 This
results in an effective average science data rate of 512 kbps. Because high temporal
resolution is desired for the ultraviolet imaging, a fast compression processor is needed.

Different techniques should be used for the different instruments, according to the
particular characteristics of each instrument. Nevertheless, the basic objective is to
produce virtually no degradation in the science data.

9.1.11.3 Mass Storage

The mass storage is the element where the science and housekeeping data are stored
uring periods of non visibility. The mass memory must be protected against the
environment, and especially against radiation. The mass memory is placed in the umbra
of the spacecraft and thus the temperature is maintained within reasonable limits during
used 1551011 PhaSeS ' and eSpe ° ally durin S the orb ‘ting phase, when the mass memory is

To reduce the mass, power and volume, advanced technology processes and high density
3-D packaging techniques must be used. The application of miniaturisation techniques is
a must given the amount of data storage required and the spacecraft system budgets.

9.1.12 Ground Infrastructure and Operations

Cost concerns and the high demand placed on the tracking networks around the world
have led to the selection of a single ground receiving and TTC station. This means that
the spacecraft will only be tracked about 8 hours a day. The impacts on the spacecraft are

discussed in section 9.1.10 above. K drtare

A high degree of on-board autonomy (as described in section 6.3.10) can reduce the

Due 1reSOUrCeS S1 g nif] cantly with respect to past interplanetary missions.
Due to the overall mission cost constraint, however, the required amount of new
developments must be controlled carefully.

9.1.13 SAUNA Global Budget

In table 9 6 we present the breakdown for mass, power and cost for the SAUNA mission.
appendkB ed breakdown to unit leve1 ' P lease refer to the SAUNA Mission Data in

Table 9.6 SAUNA Mission Budget

SAUNA MISSION - BUDGETS

Ml

Item

Mass (kg)

mm\

Cost (M$)

A. Structure

32.00

0.00J

4.00

B. Propulsion System

30.00

5.00

C. Power System

35.61

0.20

2.75

D. Attitude & Orbit Control

26.88

71.60

6.80

E. Thermal Protection

35.00

35.00

■Kill

F. Communications

27.00

750.00

G. On-Board Computer

2.00

9.00

6.00

H. Subtotal Spacecraft Bus:

188.49

7165.80

48.30

I. Payload

17.70

18.20

13.00

LI. Subtotal Dry Mass (1)

226.81

7219.92

61.34

L2. Propellant (52.5% of wet mass)

250.68

0.63

541.00

7902.40

74.86

(1) incl. harness; (2) incl. margin
and launcher adapter

R. Launcher capacity and cost:

697.0C

iBi

60.00

Rl. Launcher mass margin:

156.0C

S. Ground Operations

v -J : i

19.00

T. SAUNA Predevelopment (SPP)

22.50

U. Subtotal Cost:

176.36

V. System Cost Margin 10%

17.64

TOTAL MASS, POWER, COST:

541.00

7902.40

194.00

Note that the wet mass of the spacecraft in this budget is around 540 kg, which surpasses
the figure of 300 kg used in the initial feasibility studies by almost a factor 2. Fortunately
we still have a considerable launcher mass margin for the C 3 needed to achieve a Venus
transfer orbit (the launcher referenced in this table is the Delta II (7925)).

With this large mass, the mission is still feasible. The transfer time to the 0.2 AU orbit
will be significantly longer (in the order of 2.5 years), however, assuming the same 02 N
ion thruster is used. This again means an increase in operations cost which has not been
accounted for above.

The selection of subsystem components is very conservative, however, making use of
existing technology rather than speculating upon the future availability of miniaturised
systems and nanotechnology (which would reduce mass). This conservatism leaves room
for considerable improvements in performance in the course of the further system design.

The SAUNA Predevelopment Programme (SPP) introduces an extra cost of US$ 25
million (including 10% margin), which is part of the reason why the bottom line cost
figure in this table is higher than the US$ 160 million quoted elsewhere in this report
The cost could be reduced by technology transfer between potential international
partnerships in the SAUNA programme. Any improvements in relevant technology in
the Near- and Mid-Term up to the planned programme kick-off (mid-2001, see section
9.1.16) would contribute to a reduction in mass, power and cost. Nevertheless the risk-
mitigating element of the SPP, especially with respect to the qualification of the ion
engine, is an indispensable part of SAUNA.

r>^. nru^. C,Qr>i &r\rc> anH Hlimanitv

9.1.14 Technological Issues

Due to the cost limitation we have sought to use available technology to the maximum

extent. Below we identify some technological enhancements that will increase the
chances of mission success:

Propulsion: A high-thrust (>0.2 N) ion engine needs to be flight qualified
with a rating of more than 10,000 hours of continuous thrust.

Power: More efficient solar arrays in terms of W/m 2 and W/kg; lower cost
longer life, and heat-resistant solar cells.

• Materials: Lubricants to avoid cold welding; heat shield materials;
structural elements able to deal with high thermal stresses; Solar cell high
temperature bonding agents.

• Thermal: Improvements in low mass radiator and heat pipe technology;

• Electronics: Radiation-hardened memories with high capacity (Gigabit
class) able to resist large doses of radiation over long time spans.

GNC: Autonomous navigation techniques; control of spacecraft with low-
thrust ion thrusters;

• Communications: Use of phased arrays for long-distance transmission;
optical communication; developments in solid state amplifiers-
deployable/inflatable antennas;

• Reliability and Safety: The impact of the operational lifetime requirement of
5 years (total) has to be assessed for all subsystems with regard to the
unusual environment encountered in orbit at 0.2 AU.

9.1.15 Policy & Legal Aspects

In the spirit of the Ra Mission Statement [chapter 1.1] we have opted not to use RTGs for
power. In this respect our mission is geopolitically neutral.

Jnnc H h0iC f e ° f a la r? veh i cle and Iaunch site brin S s with various political
considerations, we shall not dwell on those here. However the spacecraft mass is

sufficiently low to allow a wide range of optional launch vehicles.

9.1.16 Programme Timeline

The . S , A ™^ K '™ ssio " can be launched as early as 2005, depending on the availability of a

propulsion technology With a launch date of
below°° 5, ^ SAUNA P ro J ect development scheme should take the form of figure 9.11

Fig. 9.11 SAUNA Programme Timeline

The significant elements of this programme plan are the following:

• A SAUNA Predevelopment Programme (SPP) running in parallel with
Phase A and Phase B to qualify the critical technologies (with particular
focus on the ion engine) before the start of Phase C/D

• A total design and development time (phase A to launch) of 4 years

• A total programme time of 9 years plus an optional mission extension.

9.1.17 Conclusion

The SAUNA mission is feasible with a Life Cycle Cost of less than US$ 200 million. The
SAUNA spacecraft will perform scientific measurements in the near-Sun environmen
and simultaneously demonstrate long-duration survivability for missions in this region.

9.2 Solar Threat Monitoring and Early Warning Systems

This section describes the steps taken to design a Solar Threat Monitoring and Early
Warning System for the Mid-Term, based on the applications needs and opportunities

identified in section 5.2.

The thrust of this effort is thus to focus on the design of a dedicated solar threat
monitoring mission and to evaluate its commercial viability. We therefore aim to limi
ourselves to the use of existing technology and take into consideration the heritage of
proven instruments and components.

After the introduction of our study approach [section 9.2.1], we determine the customer
requirements [section 9.2.2]. Several mission options for a dedicated early warning
system are then explored [section 9.2.3]. We describe their working principle and assess
the effectiveness of the concept. Based on that, we choose an array of heliocircular
spacecraft as our preferred early warning system [section 9.2.4]. A preliminary design
analysis is outlined in section 9.2.5. Finally, possible alternatives and scientific
opportunities are pointed out in section 9.2.6.

9.2.1 Study Logic

The following study was approached with an overall logic displayed in figure 9.12.

CREATE EVALUATION BASE

DEVELOP MISSION
OPTIONS

Oplon A 1 OpdorC
Op oort B j

CONOUCT FEASI8IUTY ;
ANALYSIS

Pr*limm*ry
Mimioo 0«s*gn j

[ SOLAR EVEN TS 1 01 ^P*^ 1

PRELIMINARY

MISSION

DEFINITION

► T«cfvw*og Tr«dM |

^8* 9.12 Logical Sequence for the Study.

9.2.2 Requirements

This sub-section examines in sequence the customer requirements, the functional
requirements, and the derived functional requirements.

9.2.2.1 Customer Requirements

The potential customers of a Solar Threat Early Warning System were identified and
described in sections 4.5 and 5.2. For convenience they are listed again in table 9 7 where

their requirements are also summarised.

Table 9.7 Early Warning System- Customer Requirements.

CUSTOMER

Power grid operators

Microprocessor manufacturers

Geophysical surveyors

Civilian HF communications

Earth orbiting satellite operators (non-polar LEO )
Earth orbiting satellite operators (polar LEO)

Earth orbiting satellite operators (MEO)

Earth orbiting satellite operators (GEO) _

Non-Earth orbiting spacecraft operators _

Military shortwave communicatio ns
Military radar and HF communications
Shuttle & Space Station astronauts
Interplanetary astronauts

Type of Warning Required

Magnetic Storms

Very High
Energy
Radiation

High

Particle

fluxes

Induced

Magnetic

Fields

Min. Time
Required (h)

1-6
1-6
1-2 1
12

15 min
15 min
15 min
15 min

1 [Tedrow, 1996]

To clarify the warning categories used in table 9.7, the relation between events on the Sun
schematic of figure 5*^3* °" P ° SSib ' e CUS '° merS [SeC,i ° n 5 21 * Summarised in the

Fig. 9.13

Connection between solar phenomena and effects on the ground and on
space systems (Energetic particle emissions shown only for reference).

Nature of early warning

The nature of the information provided to the client as part of the warning should
include the following estimates:

i) time to impact,

ii) severity of impact,

iii) duration of impact.

Future work should examine the accuracy and tolerances with which the client requires
event time, event magnitude, and event duration information.

Tarcet market

A top level decision was made at this point to focus only on those clients which are not
shaded in table 9.7. This was based on an assessment of the commercial potential of the
customers. Unsurprisingly the selected clients all have systems inside the

magnetosphere.

Nature of Threat

The nature of the threat for our target commercial market is thus geomagnetic storms. Our
target product can now be described more precisely as a Geomagnetic Storm Early

Warning System.

9.2.2.2 Functional Requirements

Here we specify at levels of increasing detail what functions the Geomagnetic Storm
Early Warning System must be able to perform.

T.evel 1: GENERAL

The Geomagnetic Storm Early Warning System shall:

• notify clients of solar triggered events which threaten their systems.

• r a-TV »p Sim for Science and Humanity

• notify clients of the expected time, magnitude, and duration of impact,

• include estimates of the risk to the client's particular type of system,

• provide value added information on how the client's particular system is at
risk of being affected.

Level 2: MAGNETIC STORMS

In order to provide warnings to the operators of systems within the
magnetosphere, we have the derived functional requirement that the
Geomagnetic Storm Early Warning System shall:

• be able to predict when magnetic storms will occur,

• be able to predict the duration of the magnetic storm,

• be able to predict the intensity of the magnetic storm.

Level 3: TIMING

Based on the Level 1 and 2 requirements, as well as the customer
requirements, we can identify more specific requirements, i.e. the
Geomagnetic Storm Early Warning System shall:

• be able to detect the triggering phenomenon of a magnetic storm at least 12
hours prior to storm initiation,

• be able to forecast the onset time, such that it will happen during a 90
minute alert period starting at the specified time.

Level 4: PHYSICS

Since geomagnetic storms are thought to have numerous triggering
mechanisms (see the physics background of section 4.3) the early warning
system must be able to detect all of these. Thus, the Geomagnetic Storm
Early Warning System shall be able to detect:

r, aVeS ( such as those which result from Coronal Mass Ejections
(CMEs) like magnetic clouds, and Corotating Interaction Regions (CIRs)
caused by high speed solar wind) which threaten to impact the Earth's
magnetosphere [Chen, 1996] [Farrugia, 1996] [Green, 1996],

• interplanetary magnetic fields (IMFs), with a large intensity and long
duration southward component which threaten to impact Earth's

magnetosphere [Gonzalez, 1996].

The two phenomena above will directly dictate the minimal instrumentation chosen in
the scenario described below. Note that both phenomena are thought to have at their
root a solar event of some kind. In particular, they have been found to occur often in the
presence of, or after the occurrence of:

• a solar flare and

• a radio emission burst .

In the future, given a sufficiently accurate model, it may be sufficient only to witness the
original triggering event at the surface of the Sun and compute (with knowledge of the
state of the magnetosphere) whether or not a magnetic storm will result, and if so: when
for how long, and how strong. For a Geomagnetic Storm Early Warning System using
current state-of-the-art models it is felt that this is not realisable in the Near or Mid-Term.

Mid-Term Programme •

Nonetheless, performance of the Early Warning System would likely be enhanced by
measurement of the above solar events.

Level 5: TRAJECTORIES

In order to be able to predict impact of the phenomena described in the Level
4 requirements it is necessary that the Geomagnetic Storm Early Warning
System be able to:

• predict the trajectory and evolution of interplanetary shock waves,

• predict the trajectory of southward interplanetary magnetic fields.

This results in the derived requirement that the system be able to:

• measure the position and velocity of the given phenomenon.

9.2.3 Magnetic Storm Early Warning Operational Concepts

The Level 3 and Level 4 requirements that we introduced in the previous section imply
that we need to detect triggering mechanisms for geomagnetic storms, le. shock waves
and dangerous IMF's, well in advance, before they hit the Earth.

For that we envisioned several physical methods summarised below.

Possibilities to
detect

interplanetary
plasma structures

\

/

remote sensing

in situ:

passive:

active:

magnetometers
plasma analysers

Neutral Atomimaging
Thompson scattering
ground based radio arrays

Radio Plasma I maging
Faraday rotation

Figure: Physical methods to detect DIPS

To localise plasma inhomogeneites a variety of methods can be used, like Neutral Atom
Imagine [Imager for Magnetopause-to-Aurora Global Exploration, WWW] and in situ
plasma analysing [Mars '96 FONEMA, WWW. All of these methods will be introduced
and evaluated in the different mission concepts we present in this section. However, we
want to stress already now, that one needs to use in situ measurements to measure the
strength and direction of the interplanetary magnetic field.

Several mission concepts for a dedicated early warning system are briefly explored in this
section. Later, they are judged [section 9.2.4] based on their expected fulfillment of the
requirements. From this assessment, an array of heliocircular spacecraft is chosen for a
preliminary design analysis [section 9.2.5] and some alternatives for further study are
identified [section 9.2.7].

9.2.3.1 Option A- Heliocircular Array of Spacecraft

Mission Description

This mission consists of sending a fleet of (small) satellites into an orbit around the Sun
(in the ecliptic plane), performing in situ measurements, as shown in figure 9.14.

• Rr The Sun for Science and Humanity

Fig. 9.14 Orbital configuration of option A.

Working Principle

Equipped with magnetometers and plasma analysers, this system will be capable of in
^measurements of both interplanetary shock waves, and southward interplanetary magnetic

mul^bedpl 0 /^ 6 S H aCe r aft S f h ° ulc ! be de L nS6 ' 80 that CMEs and magnetic clouds

could be detected and information about their properties forwarded to Earth.

9.2.3.2 Option B- Indirect Sensing via Spacecraft at L4/L5

Mission Description:

Two spacecraft at Lagrangian Points L4 and L5 send pulsed radio signals
and analyse them. Measurements are then forwarded to Earth, as shown in

to each other
figure 9.15.

Fig. 9.15 Orbital configuration of option B.

Working Principle:

In order to give warning of the most serious single cause of geomagnetic storms - large
scale (prolonged) strong southward magnetic fields - the interplanetary magnetic fiefd
could be sensed by the Faraday rotatlon induced in transmitted signals welUbove the
plasma frequency. In addition, some measure of the average density could be gained
from a measurement of the signal loss due to scintillation. Previous studies have

rcreen e i 9 g/ adl0 S °“ ndl " g of solar wind on smaller scales near the magnetosphere
[Green, 1996 proposal] and transmission-probing of the solar corona (with a much higher
plasma frequency) from an anti-Earth orbit [Patzold et al, 1996]. 8

Mid-Term Programme • 233

Preliminary Analysis:

The Faraday rotation angle <|>, by which the linear polarisation of a transmitted radio
wave at frequency (0 is rotated, is [Benz, 1993]

In e'

m] c 2 (O 2

\ n ‘

B cos 6 ds

m e l uj -

where the integration is carried out over the viewing path length and the factor B cosd
sees only the magnetic field component parallel to the viewing path. This poses a coup
of problems for the remote detection of magnetic cloud-like structures. First having
spiral configuration, the strong field of a perpendicularly-oriented magnetic cloud would
average tolero in the line integral. For the case of a magnetic cloud whose symmetry
axis is lying in the ecliptic plane and perpendicular to the Earth-Sun line (this case has the
highest^southward magnetic field impacting the geomagnetosphere) there would be a
net Faraday rotation, but the effective (parallel) field strength woul ^ be ^ u 7 C ^^ SS so th t f l , "
that of the true magnitude. Unfortunately, the L4-L5 distance is about 1.7 AU, so the
summed effect of many smaller-scale field variations could overwhelm the signal from a
magnetic cloudeven with diameter 0.2 AU, suggesting that this technique be put to use
on a smaller scale. Still, assuming average magnetic cloud parameters from [Lepping e
al, 1990], a Faraday measurement with signals of 30-50 MHz could give use u warning

information.

Communication Considerations

Difficulties of this proposed system are required antenna size, power demand and
information content of the weakened / refracted radio waves. Compressed pulse
techniques similar to those used in radar should be investigated to support this option.

9 2.3.3 Option C- Solar Wind Event Imaging and Tracking (SWEIT)

Mission Description

The SWEIT (pronounced "sweet") Early Warning mission uses a combination of new
kinds of imagers to detect Interplanetary Plasma Structures (IPS) which emanate from the
Sun and threaten Earth satellites and Earth systems. In addition, it provides simple white
light imaging of the upstream limb of the Sun.

The mission uses two identical spacecraft, one located at L4 and the other at L5, in order
to provide a 3-D imaging and tracking capability, as shown in figure 9.16.

L5

Fig. 9.16 Orbital configuration of option C..

234 • Ra: The Sun for Science and Humanity

Working Prinriplp

corona is not discernible against the background of the Sun from

For effective remote sensing of the IPS, two spacecraft provide a stereo view Thp
possible means of imaging are discussed in section 9.3.3.

Preliminary Analysis

Neutral hydrogen in the energy range of lO'-lO 3 keV [section 9.3.3] travels no faster than
for eaHv d ' makmg 9 n *; Utral Partide imager (NPI) with a hi S her energy range necessary

5ST f0t USi " 8 radi ° S ° Und,n8 “ “ interplanetary tZl >oul/Zl

9.2.4 Trade-Off of Solar Warning Missions

are -« — •
Table 9.8 Early Warning System qualitative trade-off matrix.

Performance:

+ good ofair -poor ? unknown

Based on the above trade-off, scenario A (the Heliocircular Arrav nf <;n^ flA

Detailed assessments and further trade-offs for the Heliocircular Spacecraft Array class of
mission are described in the following sub-sections.

9.2.5

Preliminary Design of Heliocircular Spacecraft Array Concept

Pavload Requirement Estimates

instruments are listed in table 9.9.

Table 9.9 Payload Estimates for Heliocircular Array Spacecraft.

Mass (k«)

Avr. Power (W)

Data Rate (kbps)

Comments

nbcma analvser

r 6.0

i 4.0

1.2

aiiwiT

magnetometer

3.3

1.9

0.5

Including boom

Communication Conside rations

Th e -TJTeTufan^I ^

SZ§£ Tta co—a n Hon y s architecture will have to deal with *ese proWems

offs and considerations is carried out in section 9.2.5.2.

Selection of Orbital Radius and Number of Spacecraft

The factors driving the number of spacecraft required follow from the requirement of
geoma^etks^oms^ n ^owingthe a ^ze^Hhese°fe 1 amres S an^hdrjiropagatk)n speed, one

and outside or u al, 1990], approximately 20 spacecraft are needed

Independent onheir^ohJr'nfdius to'ensuri "complete coverage” - i* .ha, each cloud is
detected at least once.

Tthe ecliptic or are only slightly inclined, and thus the relevant cross-sechon would be
much larger than 0.28 AU.

For lack of better understanding, we take the size and evolutionary behaviour of
magnetic clouds to be representative of CMEs in general. The heliosphenc array should
also^ive ample warning of CIR-associated shocks, since these can be inferred from both

the location of fast- and slow-moving solar wind regions, whose counterparts
on the Sun are generally long-lived, and

the location of the CIRs themselves: since they are corotating, they would
often be sensed by several spacecraft in the array before reaching the Earth's
solar longitude.

Optimization of Spacecraft Solar Radius

Fig. 9.17

Optimisation of heliocentric distance. Several parameters considered in
the °P llI ? Isatl0n °f orbital solar radius for the spacecraft arrav. The "total
cost is the gross wet mass of the spacecraft fleet. The planetary
perturbations are due to Mercury, Venus, and Earth.

Based on the use of solar electric propulsion and the Av's required for various circular
solar orbits, candidate wet masses were calculated [table 9.12]. Some Examples of
calculated Av values are listed in table 9.10. r

Table 9.10 Av values for several heliocentric distances.

Distance (AU)

0.3

0.18

0.5

Av (kms 1 )

22.6

34.51

11.98

The total Earth-launch mass (plotted as "cost") of the spacecraft array is shown in figure
9.17, based on a number of spacecraft intermediate to the two extremes. Based on this
cost profile and on the degree of advance warning provided by heliocentric arrays at
different solar radii, which is shown in figure 9.18, an orbital solar radius of 0 5 AU was
chosen for study.

Mid-Term Programme • 217

Solar Event Advance Warning Time

1 20 j

^ m n in « »« ■* JO

M O ^ O

SOLAR ORBITAL RADIUS (AU)

Fig. 9.18

Solar event advance warning time. The minimum advance warning time
for arrays at different radii results from the length of solar conjunction at
that radius, and the time of propagation of fast solar wind structures to
Earth after being sensed.

Mass requirements

Mass requirements for a single spacecraft from the heliocircular array were estimated as
seen on table 9 11. The payload and communications hardware mass were determined
from the equipment described earlier in this section. A dry mass of 55 kg was then
estimated from these values, based on general historical trends for small spacecraft
[Larson 19961. This mass estimate was then used to approximate the values for the res
of the subsystems. A total spacecraft mass of 159 kg was then obtained by adding
propellant and propulsion hardware mass estimates to the estimated dry mass. These
requirements provide only a general idea of the mass that may be required for a single
spacecraft. Further study will be needed to obtain a greater degree of confidence in the

mass estimates.

Table 9.11 Spacecraft mass distribution.

Spacecraft Subsystem

Mass

(kg)

Dry Mass
(%)

Payload

9.3

17

Structures and Mechanisms

11.0

20

Thermal Protection

2.2

4

Power

16.5

30

Communications

10.0

18

Guidance, Navigation and Control

3.3

6

Propulsion (RCS)

2.8

5

Dry Mass

55.1

100

Propellant Mass

59.1

Propulsion Hardware

44.5

Total Mass

158.7

Table 9.12 Total mass launched vs. distance from the Sun.

Distance (All)

0.1

0.2

0.3

0.4

0.5

0.8

Mass Launched (kg)

1578.1

527.5

266.9

216.1

158.7

117.6

9.2.5.1 Communications Concept

The baseline for the communications is that only those spacecraft located in the arc of
scientific interest will need to transmit their data. Given the constellation's distance from
the Sun, on-board electronics are not an issue and can be used to reduce the transmitted
data to a simple warning signal, together with some parameters characterising the
phenomenon. It will significantly reduce the data rate. To cope with the solar

conjunction problems, some geometrical analyses have been conducted in the following
section. 6

Solar Conjunction

The geometry of the link is represented in figure 9.19, showing the solar conjunction cone.

( 1 fo 9 ‘ 1 ?u th o °f S ° f Slgm due to solar con i un ction considering a Sun view of angle
of 1.5 from the Earth is approximately 21 hours. This leaves enough warning time if on¬
board storage is considered. This latter option consists of storing detected threatening
events and simply waiting for the spacecraft to exit the conjunction cone instead of using

).5 AU

Fig. 9.19 Communications link geometry.

Thermal Noise

In order to avoid the drastic increase in thermal noise due to the Sun s background
radiation, we will assume that communications are interrupted as soon as the Sun enters
the major lobe of the ground station antenna.

Antennas and Transponder

In order to implement the communications design that has been discussed, each
spacecraft will be equipped with a classic X band transponder. The advantages in the
Mid-Term time frame of this band has been assessed in section 6.9. The spacecra t
antenna will use advanced concepts such as phased array techniques that have already
been addressed.

Ground Segment

Continuous coverage is required on Earth in order to monitor any threatening solar
event Therefore, it is highly unlikely that the Deep Space Network would be available
continuously for our ground segment. Instead, we propose to explore the use of smaller
antennas (e.g. 15 m) that are more widely spread and available [section 6.9].

9.2.5.2 Spacecraft Configuration Trades
Propulsion:

Two propulsion systems were traded to assess which one would be suitable for this
particular mission. The two systems considered were chemical bipropellant and solar
electric propulsion. Solar sailing was not considered due to its relative lack of heritage as
compared with electric propulsion. Based on the dry weight for the spacecraft and t e
expected Av for the manoeuvre from 1 AU to 0.5 AU, the propellant mass was calculated
for each system. The results are in table 9.13. The additional dry mass required is the
mass added to the system if solar electric propulsion is selected. However, even with the
additional dry mass added, the significantly higher performance of the solar electric
propulsion system yields a much lower propellant mass requirement. A comparison of
the total mass launched versus target distance from the Sun is included in table 9.12.

240 • Ra: The Sun for Science and Humanity

Tab, e 9.13 Comparison ofpropellant masses to propulsion systems considered for

Isp (S)

Propellant mass req'd
for Av manoeuvre
(kg)

Additional dry mass req'd
for propulsion System
(kg)

305

2996

0

3300

59.1

44.5

Propulsion

System

Chemical

Bipropellant

Solar Electric

The s °lar electric propulsion system requires 2.5 kW of power, which is significantly

reoutrpH han he , P ° Wer required for a chemical system. However, the additional mass
in^io-n f m S0 ar arra y s ' ^ accommodate the power requirement, is probably
insignificant compared to the additional propellant mass required if a chemical
propulsion system is selected. During the preliminary design phase, there should be a
trade between I sp and power required for the electric propulsion system.

Thp r if'l Ctri f M° n HaS Signi f icantly less fli § ht herita ge than chemical propulsion
The lack of flight heritage could result in significant testing requirements and

development cost for the solar electric system. Increased flight experience with solar
development cost" s 'S">f'«ntly benefit this mission by reducing the potential

2“ Av req n ir ! d '° S ° fr ° m 1 AU '° 0 5 AU P redud « the use of chemical propulsion
nrnh P hl° Pe f nt m i? SS required from a chemical system to perform this manoeuvre
probably outweighs any potential hardware mass savings gained from using it The

propeflant mass could be reduced by using gravity assist manoeuvre to augment the
chemical propulsion system. However, this option was not considered duringlhis studv
due to time constraints. Attitude control will be provided by a monopropellant chemicil
propulsion system, which is a simple system with extensive heritage.

9.2.5.3 Environmental Disturbances

The solar environment will influence the performance and the life of the spacecraft.

The thermal control system and reliability considerations have to take into account an
increased heat flux of about 5 times the value at Earth distance.

The calculated solar photon pressure is in the order of 10 12 Pa. Over 10 years or 3x10 8 s

al C «r S 10 an insi «™ fi “ n , 1 4 - ma guitude of the accelerahon arista”'fromHr
radiation pressure can be neglected. 6

The spacecraft will be affected by the gravitational effect of Mercury and Venus The
estimates forces are in the order of 10' 5 N. y Ine

9.2.5.4 System Installation Scenario

A study should be performed to compare the cost of launching one spacecraft at a time
on a small launch vehicle versus launching more than one on a larger launch vehicle The
earliest possible launch time frame would need to take into account the t me fo design
development and testing of the spacecraft. The time and cost for design, development
and testing could be reduced if the programme is able to take advantage of the heritage
gained from vehicles with solar electric propulsion that may precede it. ^

9.2.5.5 Costing

The costing of Ra application project takes into account the technical specificity s
(previous paragraphs) and comes from the global costing study of Ra Design project

[section 9.7 and chapter 7].

Table 9.14, figure 9.20 and figure 9.21 present the Cost Break down Structure used for the
cost analysis 8 with the assumption for this project of twenty spacecraft with from 200
300 kg Total mass each one and use two launchers class Anane 5 or ATLAS II

Table 9.14 Ra Applications cost matrix.

MISSION COST Breakdown Sturcture

I LAUNCHER
] SPACESEGACNT

Fig. 9.20 Mission cost breakdown.

As a conclusion about the mission cost, the break down of it gives 14 % for the ground
segment, 27 % for the launchers and 59 % for the Space segment.

According to the conclusion of the optimisation of the payload to the launcher capability
[section 9 7], the total price of the mission is pushing down, in using only two launchers
such as Ariane 5 or ATLAS II AS. But due to the heavy mass of each spacecraft, the cost
percentage of the launcher (30 % of the total cost) is a normal value and cannot contribute

to push the cost of the mission.

SPACE SEGMENT COST Break down Struct

■Propufcon

14 %

9% 0%

| BPowar
□Structure A
Orh«mal

B&ndance, rogation A
Control

®Co m m u n c at >o n s

ttnterma«ton A Orta
Handling
□ahers (Bus)

Vnstrumfflta ons

•Co mmu r>c a >o n s

Oln formation A Data
Handling

PQhers (P^bad)

Fig. 9.21 Cost Breakdown structure of Space segment.

™t C ™ Cl “* n ° f the “f drivers stud y< f ° r ■» normal spacecraft with a mass from 200 to
push down'tie”totaTprte P ^ ”' he n ° rma ‘ leChn °W be used in order to

As a general conclusion if the global cost is around $895 million, the learning effect has to

e taken into ac «>unt, because of the manufacturing of twenty similar satellites So the
global price would have to be pushed down. ' 6

As menhoned in section 9.2.5, the space segment cost may be overrated by up to a factor
plrformance’ U “ ° VereStimale of the » f spacecraft needed teacceptab Z

9.2.6 Further Options and Recommendations

SC ° Pe and depth ° f ° Ur StUdy ' there remain a number Of possibilities
for fruitful further investigation, relating both to the heliocentric array and to other

mission ideas. These involve innovative funding arrangements, modularity of the
recCogy SyStem ' ^ alt6rnative early warnin S s^tem requiring more advanced

The heliocentric array lends itself well to modular deployment. Because the full svstem

could h t0 be P r0l ? lblhve u f0r P rivate industry, the effectiveness of a heliocentri/array
could be shown with a subset of the approximately 20 spacecraft recommended 7
nsertion of additional spacecraft into the grid over time would still allow the benefit of

lavlo^p 8 f 6 a,S ° offerin 8 the 0 P b on of changing or augmenting the new
payloads. For instance, scientific imaging instruments could be added with support from

space agencies, or radio sounders could be added based on the success of the system and
support from industry or military. y na

In feet d is of note that several of the alternatives considered in section 9 2 3 bear a
resemblance to science missions. This simply reflects the crudeness of current
understanding of solar causes and near-solar evolution of CMEs and the solar wind
Indeed, the heliocentric array considered here has its counterparts among scientific
mission proposals, such as the "String of Pearls at 0.8 AU" and the "String of Sails at 0 5
AU mentioned in [Russel, 1996]. It follows that "mixed funding" missions are a logical
compromise, and in fact putting space weather warning instruments on various inner-

Zss^iX^Zm may be m0re PraCliCal than 3 dedicated applications

Solar Parachute as Alternative to the Heliocircular Array of Spacecraft

A different approach to the concept of the proposed system was referenced in a personal
communication between Lt. Joel MCray and Capt. Randy Tedrow of the US Air Force
dated January 16, 1996. The following information concerning stationkeeping ot
spacecraft between Earth and Sun by means of solar parachutes was discussed.

Inflatable solar parachute in heliocentric orbit used for station keeping for an orbit that
maintains a constant Earth-Sun-spacecraft angle (4°). The solar pressure on the parachute
reduce the effect of the gravity force due to the Sun, which allows the spacecraft to
emulate the rotational velocity of the Earth at 1 AU from the Sun. That gives 4 hours

warning.

A plasma analyser, a magnetometer and an energetic particle detector will be used.

The spacecraft will have a total injected mass of less than 156 kg, of which instruments
comprise 10 kg. The instruments, which require approximately 10 W, produce data at

300 bps.

The spacecraft uses a 140 m diameter deployable kapton parachute and should be
stationed at 0.4° in front of the Earth, on a circular orbit at 0.9 AU, for a period of one
year It should provide a Av=2.146 kins’ 1 to achieve its final orbit. To get in such a orbit
two Venus gravity assist plus perihelion and aphelion manoeuvres has been planned.

There are three important issues that determine the feasibility of this mission. The mass
of the solar parachute must fit within launch vehicle weight limits. The attitude and orbit
control of the solar parachute must be stable. The development cost of an inflatable solar
parachute may not be excessive.

9.2.7 Conclusions

The section 'Solar Threat Monitoring and Early Warning Systems' concludes with an
assessment of a proposed mission of a heliocircular array of 20 spacecraft. Their orbit is
at 0.5 AU from the Sun, the overall operational time of the system is assumed to be 10

years.

This mission was selected amongst other generated options under incorporation of
potential customer's requirements and an assessment of impacts of threatening Solar
Events. The design was driven to a significant extent by the requirements that it be
completely applications-oriented and that it be able to give warning of the direction of the
magnetic fields impinging on the geomagnetosphere. The latter condition dictated the
use of in situ field measurements.

The chosen mission appears to satisfy most of the requirements developed in section
9 2 2 2- it provides warning of the time of onset, the intensity, and the duration o
magnetic storms caused by shock waves and southward interplanetary magnetic fields,
assuming only modest performance of prediction models. It is difficult to judge whether
the timing precision of the forecasts would be better than 90 minutes, and in the case of a
very fast-moving CME detected during a solar conjunction, the warning time could be

less than the required 12 hours.

The preliminary cost estimates for the selected mission were rather high, but may need to
be supported with deeper analysis of the mission details.

944 • Ra- The Sun for Science and Humanity

9.3 Solar Stereo Mission

9.3.1 Introduction: Trends in Space Science Instrumentation

The rapid pace at which technological advance is affecting the menu and specifications of
spacecraft instrumentation is extreme when compared with the normal lifetime of
mission planning and design. Particle instruments are able to measure full hemisphere
vector fluxes with rapidly increasing energy, mass, and temporal resolutions, while
becoming ever smaller.; Vector magnetometers maximising the use of VLSI technology
have achieved the size of coins. The upper energy bounds of hard X-ray and gamma-ray
imagers change steadily, as demonstrated by the Fourier-transform imaging
instruments on SOHO and on the proposed HESI mission.

As the spatial resolution of remote-sensing instruments improves dramatically the
instruments are increasingly termed "imagers", and such imagers now exist for energies
from infra-red to gamma-ray. For in situ instruments such as magnetometers and
electrometers, the complexity of interplanetary and magnetospheric plasma interactions
has made multiple spacecraft increasingly desirable, as determined by physical spatial
scales larger than those of a single spacecraft. The recently-attempted Cluster mission
and the study and use of "picosatellite" swarms are examples of this trend.

9.3.2 Advantages of a Solar Stereo Remote Sensing Mission

Just as spatial structures cannot be adequately deconvolved in interplanetary and
planetary space by the one-dimensional sampling afforded by a single spacecraft it is
increasingly evident that the critical structures near the Sun cannot be understood with
two-dimensional models or a two-dimensional view afforded by a single imager Thus
the concept of Sun-observing spacecraft well away from the Sun-Earth line is the next
step m the progression towards high-resolution, 3-D remote and in situ sensing The
major advantages of such a mission are outlined below.

1. Although some of the spatial scales likely to be important in coronal

dynamics are too small for remote observation [Emslie, private
communication, 1996], the use of stereo observations in the EUV and X-ray
energy regimes are virtually crucial for resolving the 3-D structures
responsible for coronal heating and solar wind acceleration. This primary
aspect of a stereo mission has been elaborated on elsewhere [NOAA 19961
[Solar Stereo Mission, WWW]. ' J '

2. Placing a spacecraft well away from the Sun-Earth line gives another spatial
data point for in situ plasma measurements, in the spirit of picosatellite
arrays, of Cluster, and of the ISTP satellites altogether (see above).

3. As well as providing the obvious opportunity for observation of the Sun
surface over a wider range of longitude, and at a longitude more directly
affecting the Earth's magnetosphere (due to solar rotation), having an
observatory out of the Earth-Sun line presents an ideal opportunity for

viewing" the interplanetary space through which the Sun affects the Earth

^cmT SSi ° nS (f ° r , m , StanCe ' U1 y sses ' Yohkoh, SOHO) and entire campaigns
(STSP) have recently focussed on the Sun's interior and near corona and on
the near-Earth "geospace"; however, very little is known about the large-
scale structures or the propagation of CIR's, CME's, and magnetic clouds in
the interplanetary space. These interplanetary plasma structures (IPS)
undergo great evolution in between where they are measured remotely
through X-ray imaging on the Sun and where they are felt in situ as single-

D,

point solar wind measurements near Earth or as magnetometer and other
measurements inside the magnetosphere and on Earth's surface.
Contributing to this poor coverage are the extremely small densities (less
than 100 protons per cubic centimetre at 1 AU) of IPS outside about 30 R S/
and the extreme nature of the Sun's emissions, which prevent imaging of the
Sun-Earth interplanetary space from Earth or from near-Earth orbit.
Nevertheless, several methods for imaging such tenuous structures exist
(section 9.3.3) and if placed on a remote "stereo" spacecraft, could (1)
profoundly influence our understanding and predictive abilities of these
propagating phenomena, and (2) give us up to a few days of warning when
an energetic region is destined for the Earth.

4. For the previous three reasons, a solar stereo mission is an ideal candidate for
industrial involvement. Even without increased predictive power, the ability
to see coronal emissions from between the centre face and the east limb (as
seen from Earth) of the Sun, to see the evolution and trajectories of CME's
and other IPS propagating towards the Earth, and to measure the magnetic
field and particle signatures of corotating structures before they reach the
Earth are all very useful for providing warning of impending space weather
storms at Earth. This is important for several reasons: (1) such an application
is another primary objective of the Ra study [section 5.2]; (2) given the
various planned megaprojects of orbital comsat arrays for the very near
future, this early warning information will have a large and increasing
commercial value; and (3) involving industry in the planning and financing
of such a mission is an excellent paradigm for new trends in science funding,
given contemporary fiscal constraints.

5. Having two spacecraft giving stereo observations will provide technical
experience, incentive, and a baseline of orbital hardware needed for future
tomography. Tomography, which gives a fuller 3-D reconstruction than a
simple stereo view, is generally believed to require a minimum of four
separate views [Marsden, 1996]. A proposal of such an array initially may be
financially unrealistic, and would likely only be cut back to a stereo mission.
Rather, making use of SOHO (or possibly its descendents) and adding
spacecraft gradually as experience increases will be most effective and
financially sound.

Thus putting spacecraft into orbits away from the Sun-Earth line for stereo imaging of
interplanetary and solar structures is an effective way of addressing both primary Ra
objectives, scientific and practical. Indeed, stereoscopy of the Sun's corona was among
the top priorities of the solar physics researchers contacted for Ra [section 5.1].

The concept of such stereo viewing has been discussed for at least 20 years, and there are
a number of proposals made recently for such a scenario, based mostly on scientific
objectives [STEREO Mission Workshop, 1996][Dere, 1996]. In order to demonstrate the
feasibility of such a mission and of the innovative use of industry support, an entirely
applications-based stereo mission is briefly considered in section 9.2.

9.3.3 Imaging of Interplanetary Structures

Because of the tenuous nature of the interplanetary medium even amidst CME's, most
normal photonic imaging techniques are not suitable for observing plasma structures in
the solar wind. However, sensitive UV imaging and a new technique being applied
already [Pippi Instrument Description, WWW] to denser structures (for example the
Earth's magnetosphere) containing atomic hydrogen- energetic neutral atom imaging-

heliosphere S6fUl ^ ° btaining invaluable large-scale remote sensing images of the inner

Neutral Atom Imaging

f [ for ie exa e mnlp 19 rMF X ? mined P ° SSibiHty of resolvin g CIRs, energetic solar particles
(for example CMEs), anomalous cosmic rays, and the background quiet-time

in erplanetary (QTIP) ions, by detecting neutral (hydrogen) atoms created in charge
exchange recombination between 10-10 3 keV protons and drifting local interstellar
neutrals, whose density is concentrated in the near-Sun gravity well. This study
concluded that if viewed from near 1 AU, y

• the structure and evolution of CIRs and of energetic solar particles could be
discerned,

• QTIP ions would not be discernible,

• anomalous cosmic ray ions (and thus heliopause structures) would dominate
outside 5 AU.

In addition, the energy distribution of these particles would elucidate the dynamics of the
respective source phenomena. Based on the proton and photon rejection rates needed for
such an instrument, [Hsieh et al, 1992] conclude that it could be built now.

Plans for implementing such an instrument, alongside instruments to image the solar

^ E / rth w ma S netos P here ' ar e being made as of this writing (Green et al.,
1996] [Imager for Magnetopause-to-Aurora Global Exploration, WWW1 (Pinni
Instrument Description, WWW], J 1

White Light Thompson Scatter Imaging

Sensitive white light detectors can resolve sunlight that has been Thompson-scattered off
ree electrons in the solar wind. Therefore, plasma structures with increased electron
ensity can be imaged; in fact, the extent and evolution of dense CME's have been
imaged from 1970 s HELIOS data [HELIOS CME Event Video, WWW], and a new CME
imaging spacecraft using this technique has already been proposed [The Solar Mass
Ejection Imager (SMEI) Experiment, WWW], Modern CCD technology is making this a
promising method which also lends itself well to stereography.

Radio Sounding

The technique of radio sounding of plasmas consists of measuring the reflected
components of emitted pulses over a range of frequencies. The advantages of this active

from n tL? 3 - d ™ ens f ional structure and velocity structure determined
from the delay and Doppler shift of the returned signals, and that plasma density
temperature, and even magnetic field information can all be constructed from the
fluency dependence and the relative response of two radio modes, X and O [Reiff et al,

I!" a S , wtUnrlT b6 T PI U° “ S f SinCe 1962 f0r magnetospheric plasmas (Reiff

et al, WWW] [Imager for Magnetopause-to-Aurora Global Exploration, WWW] but for

interplanetary plasmas, it requires a lot larger antennas and power levels. Nevertheless
afwSfb S ° 31 Wmd imagmS With radio soundin g has b een recently proposed [Green et

Interplanetary Scintillation

Although impractical for space-based platforms, interplanetary scintillation is mentioned
here as a useful complement to spacecraft imaging techniques. Scintillation of radio
signals from narrow sources travelling through heliospheric density fluctuations can be
resolved by arrays of radio telescopes ["Heliospheric tomography...", WWW], Because a
velocity of the plasma structures relative to the source is needed to measure differential
scintillation, these surveys are generally averaged over a complete solar rotation;
however, the resulting density maps have been shown to correlate well with active X-ray
regions in the corona [Hick et al, 1996].

9.3.3 Recommendations for Future Missions

In light of the compelling motives and promising technologies for a stereo mission to
image both the Sun and interplanetary space, such an undertaking forms an important
part of both the solar-heliospheric science community's goals and the Ra Strategic

Framework.

As discussed in section 9.4, the SOHO spacecraft is proving to be an invaluable platform
for solar science. Placing a similar (but more modern) set of instruments, augmented by
some interplanetary plasma imagers, at the L4 or L5 points is an obvious opportunity to
efficiently deploy an initial stereo system. The extremely accurate launch of SOHO has
left the spacecraft with enough fuel for 20 years of station-keeping [Huber, persona
communication, 1996], so that with some extended funding it could be kept active well
past its projected shutdown in 2004. This provides a motive and constraint for quickly
launching a newer, cheaper, but similar system into a complementary orbit.

9.4 New Heliospheric Observing Platform

SOHO (The Solar and Heliospheric Observatory) was launched in 1995 and placed in an
orbit around the Earth-Sun LI point (see Appendix A). It is equipped with 12
experiments to use the advantageous position directly in the solar wind for examination
of the medium itself and for direct observation of the Sun and its corona at several

wavelengths.

After almost one year lifetime now the spacecraft has proved to be one of our most
powerful tools for investigating the Sun. In this time a great deal of exciting data has been
returned, and over the next few years many new discoveries are likely to be made as the
data is fully analysed.

However the lifetime of SOHO is expected to end in the year 2004. It is evident that to
have at least one spacecraft at LI is useful. With a replacement of SOHO carrying more
advanced (yet smaller and cheaper) scientific instruments, this option should be one of
the main issues to be considered in the Mid-Term Framework.

9.5 The Fire Mission

In this section we give a short description of the planned Russian-American solar probe
mission called "Fire". We have also tried to understand how this mission fits into our
"Strategic Framework" and which recommendations we can offer for both Russian and
American sections.

O/IQ *

r vUa Qnn for ^ripnre and Humanity

9.5.1 Brief Description of the Fire Mission

is k a P art of th « i° inl Russian-American Project "Fire and Ice" [Vaisberg Tsurutani
1995] which .s a,med at studies of extremes of the solar system and consists § two major

• Fire: Sun flyby with Jupiter gravity assist of US Solar Probe
Plamya spacecraft.

and Russian

• Ice: flight to Pluto and Charon.

The general goal of the Fire mission probes
includes investigations of:

10 U 1C

ol uu^ ui uie excenaea

• coronal heating mechanisms and transport

• acceleration of the solar wind.

Jomna^rurh^P TT ^ With ° Ut knowled S e of the 3-D model of the solar

corona structure. Such a model is necessary to define the global context for the various

local measurements. A 3-D model of the solar corona can only be constructed through

observations of corona images at the limb, using a white-light externally occulted

coronograph during the probe perihelion part of its orbit. Therefore such an imaging

experiment is an important component of the probe observational programme.

The solar disk observations will provide information on underlying solar regions which

s S Jcrnm7Th 7 6 ? pa !! al COnneCKOn of ,he situ -easurements widflow Coronal

structures. The wide-angle observations from the Plamya will give the global structure of

Z S ° lar Pr ° be WU1 paSS - ^ ^ corona abound

the limb is usually observed as projected onto meridianal cross-sections both from the

ground at total solar eclipses and from spacecraft in near Earth orbits. Of particular
scienbhc interest, especially for the theory of the shape of the global corona, iJto obtain

m I t : h8ht COr ° n f.! maSeS j” P ro i ection °nto the ecliptic plane. The Fire mission will give
s a unique possibility to observe the solar corona from over the solar poles.

The scientific value Fire mission measurements will significantly increase due to

the^nrnhT ^ if" ^^rements and remote sensing observations of the Sun both by
the probes and by ground and space based observatories. Such measurements must
allow us to study the mechanisms of coronal heating and acceleration of the solar wind
as well as to study the solar surface, including the insufficiently exploit polarreglom

To meet the mission objectives the heliocentric orbits of the two spacecraft will have a

tefthe C A natl ° n W fri eC u° the / diptiC Pl3ne and Periheli ° n distances'of about 4 1^
for the American Solar Probe and about 15-20 R s for the Russian Solar Probe The

p ass a ges of the perihelion region for the two spacecraft are preferably simultaneous
within about one hour. Before reaching perihelion, the Russian Plamya spacecraft will
scan he solar surface with the forestalling with respect to the American Solar Probe and
with the lagging after passing the perihelion.

9.5.2 Political Considerations for Fire

POliHCal imp ! icati0ns for solar and heliospheric science surrounding
Thp S p rA Probe/Plam y a mission that must be considered in any evaluation of the projecf
The Fire mission concept has gained attention at the level of the Gore-Chernomyrdin
Conference (GCC), a biannual meeting on science and technology issues between^U S
Vice-President Albert Gore and Russia Prime Minister Viktor Chernomyrdin. At present.

such high-level attention has increased political will for the project in both NASA an
RSA. However, as we have learned from past international collaborations, pohtica
attention can be both good and bad for a project and its successors.

Because of the GCC attention, the political awareness surrounding any subsequent
mission developments for Fire will be high; exactly what that awareness brings is difficult
to predict However, an awareness of a particular tendency of political systems can e p
the space science community make its ultimate determination concerning the desirability

of the Fire mission.

Political systems have a great amount of momentum. When a political body makes a
decision it creates a precedent for itself in subsequent actions. The danger therein is that
both the reasons for the decision and the decision itself are rarely carried forward together in
the political memory. The political attention that Fire has received means that the
organisational memory for the effort is high, but the nature of that . or f* niS * b ° n *!
memory remains unaltered. If or when Fire is funded for flight, cancelled altogether, or
somehow changed into something else, there is a risk that decision brings for future
efforts in solar and heliospheric space science (whatever that decision is). For examp e, 1
Fire is funded for flight there is a chance that future requests for solar mission mig
receive a negative response because of institutional reasoning such as, "we ve already
funded a solfr mission, why another?" Conversely, if Fire is not funded for flight (even in
the case where Fire is not funded because it is felt that the money would be better spent
elsewhere in solar science), then a negative response to future initiatives could still follow
with the reasoning as, "We have already told you no for one solar mission, why bother us
with another?" Thus the decision itself and the reasons behmd it may become decoupled
and this situation poses a risk to future solar science efforts, dependent on exactly what
political bodies recall, and the mind set that they form with those memories.

The relative risks of each of these two scenarios is beyond the predictive capability of the
Ra team What the team believes is important is that a high level of awareness is
maintained concerning the long-term political implications inherent in any programmatic
decisions. This is especially relevant when the programme has gained high levels ot
political attention, as is the case with Fire. These risks must be managed if an
international framework for solar exploration and applications is to be successful.

9.5.3 Current State of the FIRE project.

The major open issue is the immediate need for RSA funding for the new generation
spacecraft to be built in Russia. Possible long lead times associated with pro]ect
implementation within Russia should be considered in light of a required project start m
1997 (necessary to meet the launch year objective of 2001). Should the impetus for a 2000
launch resurface within the Russian space community, early funding will become even
more of an imperative. Co-operative design-integration elements lack detail. Interface
definition and the establishment of technical responsibilities are required. The basis for a
successful mission relies upon the early identification of technical personnel and their
respective counterparts on both sides. Every effort should be made to facilitate the early
establishment of the necessary interfaces, protocols, and personnel.

U.S. electronic piece parts may be provided for the new Russian spacecraft. The method
by which these parts are provided is still unclear. Those parts required need to be
identified as soon as possible for evaluation of long lead U.S. procurements.

NASA funding for technology and science instrument development for the U.S. Fire
spacecraft needs to be addressed and funding sources identified. Key technology items

250 • Ra: The Sun for Science and Humanity

tn d f i mpl mc n A tati0n schedules ha ve been developed with associated costs There is a
tentative NASA commitment to fund the instrument technology development.

<« «p"- “

The launch approval process required for the FIRE mission needs further definition If
the U.S were to choose the non-nuclear power option, it is unclear what U S
responsibilities exist with respect to compliance with national and international

wSrei's°theR Ven tha ' ', he U S ' Par ‘ ° f ,he payload wil1 not contain nuclear materia
whereas the Russian part contains an RTG.

Possibdities for participation by additional countries should be explored as a method of

fCNFSU 1118 ' e> ? m P le ' there ma y be interest on the part of the French Space Agency
(CNES) to produce the thermal shield required on the U.S. spacecraft in return for French

P , art, ' ,pa , tlon , in the scientific payload. Opportunities for participation should be
development ? ' “ Peda " y When ' h ° Se arMS may COntt * ute technology

9.5.4 FIRE and Ra

t0 pe , rf0rm the FIRE mission des P ite a11 kinds of problems
ff eC "° °^ lca 7 pinanoa l and political ones. It is essential to have both of the
spacecraft in this mission since the co-ordinated measurements from the two spacecraft
are one of the main ideas of the Fire mission. With only one spacecraft the mission wi
become much less valuable. The launch by Proton launcher is the most logical way o
perform the beginning stage of the mission from the engineering point of view The
governments of Russia and, first of all, USA, must find a way to fund the mission and

launched ^ SpaC6Craft to S ether with a Russian probe with RTG aboard a Russian

sYiinimt d efmn°JT haSiS ! ? at 5* ^ miSSi ° n is SUpP ° Sed to meet ver y im Portant but

still hmited range of scientific objectives. Hence it must be just one of missions and
efforts on solar exp oration described in the Strategic Framework. The scientific results of
the Fire rmssion will be complemented by results of other Near- and Mid-Term research,
hey will also serve as a basis for future long-term missions.

9.6 Mission to Determine Biological Radiation Effects

The current knowledge relating to the effects of solar radiation has been discussed in
sec on 4.5.2. It is recognised that the solar radiation could present a significant problem
to he engineers of a future manned Mars mission. Until more information cm the effects

he r , ad ;f° n ° n the human bod y and the effectiveness of shielding is known a future
manned Mars mission may not be possible. Secondary radiation, resulting from
shielding, is potentially extremely hazardous because of high biological activity.

To quantify the radiation and shielding effects on a biological system (e.g. humans
plants, regenerative life support systems etc.), a tissue equivalent dosimetefshould be'
flown on a spacecraft. This has already been done in LEO on-board the US Space Shuttle
missions STS-60 STS-63 and STS-71. Longer duration experiments have been performed
on the Russian Mir space station e.g. DOSIMIR 1, ADLET 1, ADLET 2 and ADLET 3
These test have produced useful information [Vana, 1996], but for a mission to Mars or

the Moon the effects of radiation on biological systems above the Van Allen belts needs to
be investigated. Therefore an experiment should be flown on a spacecraft in GEO. 1 is
experiment would need to:

• measure the direct radiation,

• measure the effects of the direct radiation,

• quantify the usefulness of shielding, by measuring the reduction in
radiation, and

• measure the effects of secondary radiation resulting from the shielding.

To study the acute early radiation effects, the biological samples would only need to be
returned to Earth if methods of remotely analysing the data were not available, y
developing the instrumentation the results could be numerically coded and relayed to t
ground in the spacecraft telemetry. This would allow for the experiment to be flown as a
payload on virtually any GEO platform. It would also be interesting to study the results

along side solar event predictions.

The biological samples would need to be returned to Earth to study the delayed radiation
effects It is proposed that the samples are regularly monitored, for as long as 20 years
after the samples are returned to Earth, to determine these delayed hazards. The
challenge behind this mission is the data retrieval. One possible way of retrieving the
sample would be to fly the experiment on a spacecraft in GEO, then to return e
spacecraft to LEO were it could be collected by the US Space Shuttle. The samples could
then be returned to the ground for further analysis.

Another useful mission would be to fly radiation experiments to determine a more
accurate model of the radiation environment. The mam factors to be determined are the
temporal and spectral classifications, above the Van Allen belts. With this knowledge
more accurate simulations could be performed on Earth, alleviating the need for the
complex "return sample" missions.

An instrument that could be used in the above missions is a Tissue Equivalent Particle
Chamber (TEPC) [Margit, 1996]. This is based on an ionisation chamber filled with tissue
equivalent gas. Particles passing through the detector (mainly with cylindrical shape)
ionise the gas and produce electrons and ions. The electrons and ions are collected (for
this purpose a high voltage is necessary). The signal obtained is proportional to the
energy deposited in the detector volume by the traversing particle. The mam feature of a
TEPC is that the volume simulated by the detector is in the range of microns. By variation
of the pressure of the tissue equivalent gas the sensitive volume can be varied. It is
therefore possible to measure the energy deposition in a volume similar to tha of a cell.
Using this device changes in the composition of the particle spectrum, mainly during
solar flares, can be recorded.

An example of how the TEPC could be used on a satellite or space station is to measure
the energy deposition every 12 hours for half an hour or in shorter periods during a solar
flare. This spectrum can then be recorded on a memo card which can be analysed in the
laboratory on Earth. Another possibility is an on-line measurement. For this purpose the
recorded signals would be sent to the Earth about once per week.

The dimensions of a TEPC are approximately:

• detector: diameter 6 cm, length 10 cm, mass 0.75 kg

• high voltage supply: about 5 x 20 x 20 cm, 2 kg

95? • Ra: The Sun for Science and Humanity

• amplifier (including pre-amplifier): about 5 x 20 x 20 cm, 2 kg

• analysis device: about 5 kg

• power supply (220 V) is necessary

9.7 Mid-Term Costing

In this section we will look at costing for the Mid-Term programme of Ra. This stretches

from 2000 to 2010 and deals with the costing of the SAUNA mission and the Early
Warning System mission. J

Costing must be initiated during the conceptual and pre-development phases of a project.
It is used to determine the budget, make decisions about the future of the project
evaluate alternatives or compare estimates of the proposals. Science and the constraints
of science are increasing [Randolph, 1996]. Costing is an important part of these
constraints. It is important to minimise the costs and thereby change the public
perception about the efficiency and effectiveness of our space programmes [Scoon, 1996].

The costing tends to provide a project an iterative control process. Costing can be
divided into different phases depending on the type of project. Top level costing analysis
is used for Ra. This is suitable for future missions where factors like technology might
change. All costing in this chapter is done in 1996 US$.

9.7.1 Costing Model Used

In order to estimate the costs of the mission we use the analogy method [Wnuk 19961.
The cost break down structure is shown in figure 9.22.

Fig. 9.22 Cost breakdown structure.

9.7.2 Costing Statistics

In order to see the cost trend for solar related missions, the costs of 10 solar related
missions are analysed. Figure 9.23 shows, for these missions, the cost as a function of the
payload mass and the distance from the Sun for 10 different solar related missions. For a
detailed list of these missions see appendix A. Costing information for these can be
found in appendix D.

Mid-Term Program mp •

Kg4500 n

K$ 1.200

K$ 1,000

K$ 600

Fig. 9.23 The cost as a function of payload mass and the distance from the Sun.

The graph shows:

• For a mission located up to two hundred Earth radii, the average total cost
is around 200 k$kg ' payload.

• For a mission located between 10 and 65 Rs, the average total cost is around
400 k$kg -1 payload.

• For a mission located between 4 and 10 R s , the average total cost is around
800 kSkg' 1 payload.

For launch statistics see appendix D.

These costs are in 1996 US$. To estimate the costs for future missions the inflation must
be considered.

9.7.3 Cost Minimisation Methods

Concerning the minimisation of costs, the following global approach could be applied to
every future mission:

• Shorten cycle for conceptual assessment and design feasibility studies.,
from typically 6 months to 3 months as target for conceptual studies, and 24
months to 14 months for industrial design feasibility studies (phase A).

• Reduction of phases B, C, and D from typically 54 months to 42 months.

• Use state of the art technology or inherited space qualified hardware (no
technology development in parallel with project development phases).

• A coherent development strategy for ground and on board software,
utilising core software modules at subsystem, system and flight system
levels.

• Utilise the appropriate level of autonomy to guarantee safety, minimise
risks and reduce flight operations costs.

• Ra: The Sun for Science and Humanity

9.7.4 Cost Reduction on Ra Missions
Ground segment:

• Use common ground segments together with other missions and nations.
Better data distribution by more extensive use of Internet.

Launcher:

Use of few (combined) launchers when launching constellations,
an^thereby^helauncherco^t ^ ' he ^ Se8men ' brinSS down the ™ ss

Space segment:

" operation cos^s" 01 * 1 ^ 3 ^ 0WS Iess complex communications. This affects the

‘ v C S“^ a “^r‘ ‘° ^ laUnCh “*“" t THiS indud “ “0

9.7.5 Costing of the SAUNA Mission

In order to see the relative costs, the cost breakdown is shown in figure 9.24.

Costs [M$ ]

Oound

segmert

Laincher

Payload

Fig. 9.24 Cost breakdown for SAUNA mission.

The total cost for the SAUNA mission is $126 million i

similar as possible) missions, it is clear that SAUNA is a tow'bud'T W “ h °* ( her
launch cos, (Delta II) is similar for the

launcher cost is the cost driver The relativp met n f i *L, A snows that the
This is the cost driver on a global scale. ** f ° r SAUNA is 48 % '

Table 9.25 Cost comparison

of SAUNA with similar missions.

9 . 7.6 Costing Early Warning System Mission

The Ra Application mission consists of 20 probes,
mission is conducted to obtain some top level costing
cost breakdown for Ra Application.

A similar analysis as the SAUNA
information. Figure 9.25 shows the

Costs [M$]

500

450

400

350

300

250

200

150

100

50

0

segment

Fig. 9.26 Cost breakdown for Ra Application mission.

Maior cost driver for the 20 probes is the manufacturing of the probes. However the
Si are difficult to estimate as multiple production can decrease the cost sigmfic y.

This is due to the learning process.

t t„ now the costs for the ground segment are not estimated. One possibility is the use
U f P - t distribution company which distributes the raw data and the early warnings
concerning lolar^activity. W would mean tha, no cos, would be related ,o Ra
application for ground control.

Economic risk is bigger if only two big laun* vehicles are used rather than if many small
launchers are used, however the cost would be lower.

. . , j ¥ c nr iu p R a Application mission is on the order of $896 million. This

™fc^tS m,o m^utr P pack°a g es, starting with a constellation of 3 probe,

the cost comparison of the Ra applications mission with other missions.

Table 9.27 Cost comparison of Ra application mission with similar missions.

Chapter 10

f ^ N

Introduction

yOur View of the Sunj j

rT~>

yNeeds & Objectives I

u

■f* $ >

* <=>$tD
Technology

*01N

V^Strategi

;ic Framework

it:

Near-Term

3

if

Mid-Term

r==

\ _

Far-Temt j j

m

&*=»<=>ji | .
Political & Economic ►
Environment .A

fTpl

Conclusions ' j

Far-Term Programme

The far-term programme of the Ra Strategic Framework is designed to build on the
experience gathered during the mid-term programme. We assume that more ambitious,
lgh-cost missions are possible in the long run so long as these are balanced by a
proportionally increased economic viability, in terms of commercial exploitation and
trect benefits to society. We refer to the Strategic Framework [Chapter 2] where the
rationale for the programme is described in detail.

n the Present chapter, we will describe the various far-term initiatives in some detail,
e reader should keep in mind that these ideas were conceived and elaborated with

very relaxed constraints in the areas of funding, politics, and technology.

The far term initiatives which we have investigated are:

• Integrated Solar Science and Applications Programme

• Small Suicide Probes

• World-Wide Space Environment Forecasting System

• Preliminary Solar Power Applications

• Monitoring the Solar Constant and its Effect on Earth Climate

After the description of these initiatives, this chapter will conclude with
costing of the far term framework.

a discussion of

We urge you to approach these ideas with an open mind and to
significance in the realisation of our Mission Statement. Several of
represent logical follow-ons to the near- and mid-term programmes.

appreciate their
these ideas also

10.1 Integrated Solar Science and Applications Programme

The various missions presented in the Strategic Framework so far have been trying to
achieve a balance between the two fundamental elements of the Ra Mission Statement:
Science and Applications. In accordance with our belief that these two elements are
complementary rather than competing interests, we propose ways to enhance this
connection.

If there were funds for only one mission to the Sun, how should it be prioritised? Pure
science, or direct benefits? In the perspective of the wide open future, the very premise of
this question is unacceptable. As has been shown in this report, our future well-being
depends both on a thorough understanding of the Sun as well as the means to deal with
the threats and potential benefits of the Sun here on Earth. We therefore need a
continuing effort to address both solar science and applications.

If we remain ignorant and do nothing about the Sun and its effects, one might ask: how
much of that ignorance and passivity can we afford?

In the following, we outline ways to couple solar science and applications in an extended,
mutually beneficial far-term programme. We have considered:

• Scientific payloads "piggybacking" on applications spacecraft

• Prototype applications sensors flying on science platforms

• The use of a common bus for science and applications missions

In the ensuing discussion, we hypothesise the future reality of one or more of the
following possibilities:

□

-tf-
— *1

0

HP

r

Fig. 10.1 Conceptual Future Network of Sun-Orbiting Spacecraft.

• A fleet of spacecraft could be placed in (a) low orbits around the Sun, (b) in
a 1 AU orbit around the Sun, and/or (c) orbiting the Earth, to monitor the
variations in the Solar environment.

• These spacecraft could be used for science, applications, or both.

• The funding of such networks would be motivated mainly by commercial
applications, with spill-overs to and from science missions.

An example of a heliocentric constellation is shown in Figure 10.1.

10.1.1 Science Payloads Piggybacked on Applications Spacecraft

C ° nst f llation of s °l ar applications spacecraft is established, there are golden
opportunities for science to benefit. Not only can the data from the spacecraft senfors be
used for scientific purposes, but these spacecraft may allow for smalf and undemanding

science payloads to be accommodated, measuring phenomena not of direct interest to the
application in question. ° me

Solar scientists could get extensive amounts of science data at a low cost, and their
financial compensation may stimulate the enterprise responsible for the application

bL e rfZtawX p g ns;';:,forL elieve space science budge,s from some ° f the

There may even be a real demand from the commercial viewpoint; the potential
competition in the future "early warning" market (reliabilityand timeliness of
predictions) may drive businesses to look for new methods to derive early prediction of
solar phenomena Commercially sponsored scientific research can help providing these

new methods by looking for new relationships in solar physics.

10.1.2 Prototype Applications Sensors Flying on Science Platforms

Our next idea is just a mirrored version of the previous one. If a commercial space
venture needs to space qualify some critical new technology which would enhance their
competitive edge in, for example, the early warning market, it might look for a flight
opportunity other than its own operational vehicles. 6

Science missions could use such participation to strengthen their mission budgets, as long

as it is ensured that the added instrument does not reduce the performance of :he original
science payload. 6

10.1.3 The Use of a Common Bus for Science and Applications Missions

The realisation of a large fleet of Earth-orbiting and interplanetary spacecraft for solar
science and applications is constrained by cost. One possible way to reduce cost may be
to introduce a standardised spacecraft bus suitable for a wide range of solar missions and
Ei{ ° f ? ds , An exam Pl e o f such a standard bus might be the SAUNA spacecraft [Chapter
9.1J the design of which is not oriented towards any particular payload (however it is
constrained in the variety of missions which it can efficiently support) A serial
production of such vehicles would likely lead to a substantially lower cost per spacecraft.

A competing factor to standardisation is the desirability of optimisation of the spacecraft
bus to the needs of a particular payload. Such optimisation leads to more efficient system
performance; however, the associated development costs are higher than for a
standardised multi-purpose bus. In essence we are talking about a trade-off between
system efficiency and cost.

Depending on the scale of future solar spacecraft fleets, a balance between specialisation
and standardisation can be established by introducing, e.g., a limited number of
derivatives of a standard bus, each specialised for a certain category of payloads (e e
imaging instruments, plasma instruments, etc.) and a certain range of missions (solar
polar orbits, equatorial orbits of 0.2-0.7 AU, Earth vicinity orbits, etc.). If such a "payload-
oriented derivative bus were chosen for a mission with special needs, the penalty on
system efficiency could be brought down with respect to a multipurpose bus, while the
development costs for a mission-specific bus can be avoided.

The potential for savings through use of the common multi-purpose bus with derivatives
is a function of the scale on which such standard payloads would be flown. The success
of such a strategy would also require a strong political commitment to the continued
support of a large space fleet. In chapter 3.2, we have proposed a Working Group for
International Solar Exploration and Applications (WG ISEA), one of the tasks of which
would be to define and support the development of standard reference buses.

Furthermore, cultural changes would be needed in space communities where mission-
specific bus design is the unwritten law. Space agencies tend to support this idiom
through their desire to retain control over programmatics, resulting in an apprehensive
attitude towards externally imposed standards. These factors have effectively resisted
moves towards standardisation in the past. A recount of these cases and details on the
technical aspects of standardised buses are discussed in Chapter 6.11 of this report.

In summary, the usefulness of common buses depends on the scale of the space fleet and
the limited proliferation of special-purpose derivatives from these buses. The success o
the common bus concept is therefore a matter of economics and political priorities. The
Ra Strategic Framework strongly supports the introduction of common buses for a ong-
term programme of solar science and applications in which multiple spacecraft would fly
essentially identical missions.

10. 2 The Suicide Probe

The concept of a "suicide probe" merits attention when investigating the Sun. Such a
probe would be delivered as close as possible to the Sun or even into its surface
depending on the mission requirements and available propulsion and thermal
technology. The aim would be to relay as much data as possible before the probe
eventually succumbs to the extreme environment. Such a probe would be the first real
encounter with the Sun, which can inspire education and awareness [see Section 8.6].

We assume that a suicide probe is launched from a mother spacecraft. The mother
spacecraft will bring the probe in the proper orbit and will take care of the
communication from the probe to the Earth. From the standpoint Av, a highly elliptic
orbit of the mother spacecraft is an advantage, as we will see in the technology issues.
Options for a suicide probe are:

• A piece of appropriate material, the behaviour of the material when it
enters the Sun is observed by remote sensing instruments (on the Earth and
the mother spacecraft);

• A deceleration triggered, chemical or physical reaction (nuclear bomb), the
phenomena are observed by remote sensing;

• An instrument probe with power, thermal protection, and communication
to the mother spacecraft, which relays the data to Earth; and

• A dedicated probe with various on-board instruments, thermal protection,
power, and communication.

The choice of the probe depends on many factors, like:

• Available funding, which is closely related to the scientific community and
the public interest;

• Scientific objectives of the suicide probe; and

Constraints of the mother spacecraft, such as mass, dimensions and
launching capability.

Also the possibilities of a number of probes must be taken into consideration For the

S™ A dl c CUSS10 n V S a ? sumed that the df y of the mother spacecraft is 150 kg [see
SAUNA, Section 9.1]. The mass of the probe should be only a small fraction of the dry

mother spacecraft mass, say < 50 kg. In the next section the scientific goals and the
technological issues are discussed.

10.2.1 Scientific Goals

The scientific goals of such a probe would be to make in situ measurements in the inner
corona and deeper layers in the solar atmosphere. Plasma parameters measured in situ
would need to be combined with good contextual remote sensing observations of the
entry site (in the same way as the Galileo probe). In situ measurements would seem to be
the priority, since higher resolution remote sensing measurements may more easily be
made by improving the instrumentation's resolution than by going closer Detailed
evaluations are needed to determine whether flying closer to the Sun is really of more
beneht than investing with remote sensing technology.

10.2.2 Thermal Control

Clearly, an instrumented probe presents a huge challenge for thermal protection and
lgh-tempera ture, radiation-hardened electronics. The issues related to a heat shield are
discussed in detail in section 6.7. The maximum temperature of a plain Carbon-Carbon
heat shield (before ablation occurs) is about 3000 K. The figure below [Figure 10.2] shows
the temperature of a plain Carbon-Carbon heat shield (angle of incidence is 30°). At a
distance of less than 2 R s ablation will occur, which would result in the evaporation of the
probe. To go closer to the Sun other materials should be taken in consideration, such as
samarium oxide, hafnium carbide and tungsten.

Fig. 10.2 The Temperature of the Heat Shield Near the Sun.

10.2.3 Propulsion

The propulsion system needs to provide the Av required to cancel the orbital velocity of
the mother spacecraft from which the probe is deployed. To estimate this Av three
trajectories are analysed, from the 0.2 AU circular orbit:

Concl iicinn •

1. Directly to the Sun (Circular), by decrease of the tangential velocity

2. To a gravity assist of Mercury (Mercury)

3. Increase the aphelion (from 200 - 800 R s is about 1 AU - 4 AU) of the mother
spacecraft and launch the probe at the aphelion (perihelion is 40 R s )

The graph below [Figure 10.3] shows the Av needed, to launch the probe from the mother
spacecraft as a function of the perihelion (minimum distance from the centre of the Sun (1

-10 R s )).

>

3

a

Fig. 10.3 The Required Av as a Function of Perihelion.

It shows that only the "Aphelion 800 R s " mission has values in the order of 3 km/s for a
distance of 1 R s from the centre of the sun. The other trajectories all have high Av values.
The next graph shows the mass fraction of a mono propellant with a specific impulse of
220 m/s as function of the distance from the Sun's centre.

Fig. 10.4 The quotient Mass (propellant)/Mass (Total Initial Spacecraft) as a
Function of the Perihelion.

The figure above shows that for a mission to a Perihelion of 2 R S/ a quotient of
Mprop/Mtotal, say 0.8, is only possible for the trajectory "Aphelion 800". If the required
total mass of the probe is smaller than 50 kg, the required dry probe mass is 10 kg. The

OfA • Ra: The Sun for Science and Humanity

K Pr0P °f.'J lnstruments (magnetometer, e.g.) have a mass ot about 3 kg. Moreover
bout three of these are needed to obtain valuable scientific information. 8 Innovative
design solutions, (miniaturisation) and/or the replacement of electrical systems by
or ophcal components (nanotechnology) are needed to lower the mass of the
instruments. This makes an instrumented probe more feasible.

° P . ti0n ’ll 0 USe hi§h ' Specific im P ulse propulsion systems, such as an ion engine

dlrrlZ * g ‘ H Tl Ver ' ^ GXPeCt th3t the "technology and miniaturisation will
decrease the mass of the instrument significantly in the next 20 years. It will then be

possible to use a known, low risk, and low cost propulsion system for the suicide probe.

10.2.4 Communication

The following section will deal with preliminary design considerations for the

colZed b ‘° n T u his section will only look a. the scenario where data

collected by an mstrumental probe needs to be relayed back to a mother spacecraft.

The closeness of the approach, and the small size of the probe, prevent RF
communications altogether. The surrounding interference would simply expand the
requirement for RF power or impose ridiculous sizes on the probe. It is assumed that the

scientific objective for the mission would not be RF propagation study in the outer/inner
corona.

Optical links are therefore the only known solution. Spatial resolution of the signal from
R . lT rS J he , SOlar diSk iS achievable ' considering the maximum § expected

ICzichv 19961 deS ^ Ct [°'?' Acc uracies of < 1 microradian are currently achievable

[Czichy, 1996], A more detailed analysis of the Sun's spectrum in order to do a proper

achievable^ ^ fr6qUenCy 1S necessai T- A good signal to noise ratio would then be

IST^ ntifi V eqUi ff ment u for the link to be du P^x. The power budget will likely
dictate that the probe will not be continuously broadcasting information during the

V y ho . e , tra,e f ^ but rather dunn g the last phase of the mission (encounter). The probe
should be fully autonomous [See Section 6.10.3] in terms of telemetry, tracking and
control. Acquisition of the mother spacecraft as an open-loop system is a bit more tricky
as the sensors could be blurred by the proximity of the Sun. The probe could be

programmed with the relative position of the mother spacecraft with respect to the Sun or
other sources.

Finally, the need for simplicity of the communication system on the probe will be met bv
an increase on the complexity of the receiving end at the mother spacecraft One
approach could be to have the mother spacecraft act simply as an optical repeater. The
hght beam could be amplified using standard optical techniques, and relayed back to
Earth where coherent detection could be more easily implemented. It remains to be seen
if the overall link budget would allow for this, given that amplification at the mother
spacecraft would not be regenerative, and noise would be transmitted as well.

10.2.5 Alternatives Probes

An alternative to an instrumented probe is the use of an entirely passive probe. The
mother spacecraft would track the projectile and observe effects remotely (the projectile
could be designed to produce a known quantity of trace element into the plasma) In the
case of a close perihelion instead of direct entry, the parent craft could rendezvous with
the projectile afterwards to analyse the effects of the solar environment on the projectile's

constituent materials. However, detailed research is needed to determine the real value
of these kinds of probes.

10.2.6 Recommendations

For a suicide probe, the primary objective is the in situ measurement of the inner corona
and the deeper layers in the solar atmosphere. However, we think that other, more broad
objectives such as entertainment and education can be used to obtain funding for the
mission. Many people have dreamed for ages about an encounter with the Sun.

To reach the Sun, the probe will be "launched" from a mother spacecraft which is in a
highly elliptic orbit around the Sun. We expect that the technology of miniaturisation
will make it possible to manufacture low mass instruments (< 1 kg). This is required to
keep the total mass of the probe low.

With the present technology of Carbon-Carbon heat shield it is possible to survive down
to one solar radii from the surface of the Sun. However, to go closer to the Sun and to
decrease the evaporation of the heatshield new technology needs to be developed [See

Section 6.7].

The proximity of the approach and the small size of the probe prevent RF
communications altogether. Optical links are a possible solution. It is expected that the
technology will be available with proper power, size, and wavelength characteristics for a
laser communications link with the probe.

10.3 World-Wide Space Environment Forecasting System

When a space environment forecasting system has been established, we must look at the
way the information is distributed and applied: Our objective would be to maximise
benefits (in the widest sense of the word) to all humankind.

It is not the intention here to sketch the specific infrastructure needed. In the following,
let it suffice to characterise some of the measures which would be most effective in
fulfilling the above objective:

• The operator/owner (whether commercial or public) of the space and
ground segments has the right to ask compensation for the services he
provides; however,

• The rates for such services would be based on a "pay according to ability
system for which the coefficients are set by the WG ISEA [see chapter 3] or
the United Nations or another representative body of international politics.

• Space environment forecasting data can be compiled from a number of
different space vehicles (preferably using a common telemetry format!),
belonging to military, governments, space agencies, international
organisations, or commercial /private entities.

• The integration of the above data will be made within a forecasting model.
The reliability of the forecasts must be very good in order for the system to
gain acceptance in the general public, [see section 8.3]

• Developing countries will be able to increase their benefits from the
forecasting data through educational programmes run by the WG ISEA or
the United Nations. These courses would be aimed at teaching how to use
the forecasts to plan effective countermeasures.

n _. TT, rt C,

t* Qr-ionrp and Humanitv

• Finally: Political, social and financial interests have to converge in the long
run to produce the maximum net "benefit" possible. ° °

T he realisaHon of the above measures can only be made possible by a strong, world-wide

po itical consensus on how to share this data, as well as a sound and efficient global data
distribution network.

10.4 Preliminary Solar Power Applications

Whereas protecting Earth and its inhabitants from the threats of our violent Sun is a
priority objective in the Ra Strategic Framework, there is also a mandate to take
advantage of the enormous solar energy output which continuously comes our way.

In this section we shall briefly point to two applications with a common denominator
Energy from the Sun. The first application relates to the conversion of the Sun's radiation
into electric power; the other concerns the direct management of that radiation in terms of
heat and light for human habitats. The objective of this part of our far term programme
would be to prepare to meet the imminent global energy crisis [O'Neill, 1989]

10.4.1 Prototype Space-Based Solar Power Stations

It is well known that solar radiation can be converted into electrical energy by means of
photovoltaic cells. Applying this principle on a large scale in space would provide an
inexhaustible energy source. Unfortunately, the establishment of such an infrastructure
is at present beyond our financial means due to the cost of space access. To provide
energy to Earth at a globally significant scale, hundreds of large-scale solar power
stations would have to be constructed in orbit around the Earth.

A first crucial step towards the realisation of such a system would be to set up a scale
prototype of a solar power station. This prototype would serve as a demonstrator for
several critical technologies related to solar power production in space:

• Assembly and control of very large flexible structures in orbit

• Highly efficient photovoltaic arrays with long lifetimes

• Microwave transmission of power from orbit: power conversion, beam
characteristics, pointing accuracy, electromagnetic interference

• Integration of a space power system to existing networks on ground

The recipient of the power from this station could either be a ground station with a small
local distribution system, or it could be small spacecraft in the near-Earth region with a
high demand for power (e.g. ion engine spacecraft, geostationary communications
sateuites or Earth observation satellites using active instruments such as microwave

We shall not address the specific ways to implement a large-scale solar power
programme in this report; instead we refer to the report produced by the ISU Space Solar
Power Program [ISU 1992], v

The development of a prototype solar power station is a "second generation application"
with respect to the Ra Strategic Framework; initiation of this programme is considered an
important applications objective for the far term beyond the year 2020.

Conclusion •

10.4.2 Light and Heat Management Systems

The Sun provides heat and light in significant amounts. It would make sense to use this
radiation directly for heating and lighting of our houses instead of going through the
inefficient process of converting solar radiation to electricity and subsequently convert
electricity back to heat and light.

A heliostat (a rotating, sun-tracking mirror) on the roof of a building can channel the
solar radiation to the basement. Here, the light is stripped of its thermal components
through an infrared mirror. The thermal radiation is used to provide central heating,
warm water, etc.; the visible light is distributed to various locations in the building
through a network of optical fibres [Scoon, 1996]. The idea is sketched in Figure 10.5:

Fig. 10.5 Solar Heat and Light Distribution System for Buildings.

This system would give almost unlimited access to light and heat indoors (by careful
energy management, benefits are not limited to cloudless daytime!) which translates to
significant savings in electricity. The technology to do this is available now, but clue to
the architectural problems involved, installation of this system is generally difficult or
unfeasible in existing buildings. For this reason, we emphasise the importance of keeping
this technology in mind when designing the buildings of the future.

10.5 Monitoring the Solar Effect on the Earth Climate

Any long term monitoring of the solar effect on the Earth climate basically involves two
processes: first, monitoring the phenomenon related to the Sun that causes the Earth
climate change; and, second, monitoring the effects of that phenomenon on the Earth.

10.5.1 Monitoring the Solar Constant

One of the physical parameters related to the Sun that can be directly measured is the
solar constant. Solar constant is a measure of the amount of solar electromagnetic energy
that falls on a unit square area per unit time at the mean Earth-Sun distance (1 AU).

The current measured value of the solar constant is 1.37 kW/m (± 0.02 kW). It is
believed that this value of the solar constant is not actually constant with time. It has
been shown by computer models that "fluctuations in the solar constant exceeding a few

268 • Ra: The Sun for Science and Humanity

tenths of a percentage would have significant climatic effects" [Evans, 19851 Thus an
accurate measurement of the changes in the solar constant, which is possible only if the

Earth cUmate * ^ ^ * l0nS Peri ° d ° f time ' i§ VitaI f ° ° Ur understanding^ of the

P* mon ‘ toril lg the solar constant from the Earth, however, has not been accurate
because the Earth atmosphere absorbs any electromagnetic radiation within the
wavelength range of 2.5 pm and 0.3 pm.

Similarly, monitoring of the solar constant using space satellites, as done by the Solar
Maximum Mission, is also possible only for a short duration. This is mainly due to two

Ae E r arT™T S: P*£* ™ easur u eme " ts are skew ed by the electromagnetic noise from
the Earth, and second, by the technical challenge of keeping a satellite, like any airborne

ject, m orbit for a long duration due to the problem of running out of power-supply,

the possibility of the satellite spinning out of control, and the difficulty of doing any kind
of maintenance work. 6 y

?«° n £ aSt/ lf tHe measuring instruments were to be placed on the Moon, much of the
difficulties mentioned above could be overcome. Because of ultra-vacuum on the Moon,
un 1 e on e arth, there is no radiation absorption by the atmosphere. And, as the
distance between the Earth and the Moon is considerable, there is not much problem of
electromagnetic noise from the Earth as with satellites on the Earth orbit. Furthermore
the solid ground on the Moon also provides advantages over the Earth and the satellites-
the seismic activity on the Moon is very small as compared to the Earth, and thus
measurements are more stable. Once the instrumentation has been left in place, it can be

sateUite 6 ^ 3 ° ng Wlth ° Ut having to worr y controlling its orbit, like in the case of a

If the instrument were to be placed on the near side of the Moon, it would be able to take
fourteen days of continuous measurement of the Sun. If a second instrument were also
placed on the far side, the two would provide a continuous uninterrupted measurement.
Buh instead of placing two instruments on two sides, it would be better to place an
instrument on the North or South Pole of the Moon [Burke, 1985], That way using onlv

one instrument, it would be possible to make a continuous measurement of the solar
constant.

There are two instruments in use at present to measure the solar constant: the older Pvro-
(ACREM)**' and ^ m ° re adVanCed Actlve CavUy Radi °meter Irradiance Monitor

There are some technical issues that need to be considered for any of these instruments to
function properly. For example, the instrument should be placed in a very high support
system to protect it from lunar dust. Some shielding should also be used to protect it
rom micro-meteorites; cosmic rays, as there are no radiation belts to act as natural
shields; and finally, light from the Sun, and the light reflected by the Earth.

10.5.2 Monitoring the Earth Climate

The monitoring of the effects of the solar constant on the Earth climate is made difficult
by the fact that the fluctuation in solar constant is only one of many factors involved in
the Earth climate change. This, however, does not rule out the significance of measuring
the overall climate change on the Earth from the Moon. 8

Uke the monitoring of the solar constant, monitoring of the effects of the solar constant
on the Earth can also be done either from Earth-orbiting satellites or from the Moon.

Conclusion • 269

10.5.3 Conclusion

In the long term (beyond 2020), the continuous monitoring of the solar constant and its
effects on the Earth climate should be given a high priority. Initially this should be done
from satellites, but as technologies mature and the cost of space access goes down, a
permanent observation base at one of the Lunar Poles will give the best observing

conditions.

10.6 Costing of the Far-Term Programme

Due to the high uncertainties of future developments in global politics, economy,
technology and infrastructure, we refrain from even attempting to cost the far-term
programme for 2010-2020 and beyond. We recommend instead to take another look at the
costing of this programme after the year 2005, when lessons learned from the near-term
and many experiences from the mid-term programmes are available.

10.7 Conclusions

For the Far-Term Programme we advocate the following:

• Integrated Solar Science and Applications Programme

• Small Suicide Probes

• World-Wide Space Environment Forecasting System

• Preliminary Solar Power Applications

• Monitoring the Solar Constant and its Effect on Earth Climate

With these missions, we fulfil the following objectives:

• Reducing cost by co-operation in areas of common interest and by
exploiting free opportunities

• Exploring the acceleration and heating in the solar corona by means of in-
situ measurements

• Enhancing the benefits of a space environment forecasting system for all
humankind

• Exploring ways to solve the imminent global energy crisis on Earth

• Study the impact of variations in the solar output on the Earth s climate

The Top-Level Recommendation for the Far-Term Programme is:

• To focus mainly on the fulfilment of the application objectives related to the Sun as
both a source and a threat as well as on the fulfilment of scientific objectives related
to the solar physics and theory.

070 • Ra: The Sun for Science and Humanity

Chapter 11

Conclusion

The Ra report is a call to action. Knowledge of the Sun is vital to us as humans and to our
planet. Our star deserves our attention and study.

We, the Ra team set out to explore strategies that would increase our understanding of
the Sun and its effects, and that would help us apply solar knowledge for the benefit of

humankind. We did this through an international perspective and we document our
strategies here.

The potential for solar science and applications inspired us to question and to investigate
various issues. Strategic planning moulded our investigations into a rationale It enabled
us to formulate a programme or, as we have called it, a Strategic Framework. Policy
defined the environment in which we could organise and operate. Costing, marketing
funding, and technology all served as a check and balance, reminding us of reality.

We recommend solar explorations and applications in three time frames: a Near-Term
(1996 to 2000), a Mid-Term (2001 to 2010), and a Far-Term (2011 to 2020 and beyond).
Each technological, economic, and political issue fits into one of the time frames In the
realm of solar science and applications, deciding for this Strategic Framework means
beginning a process that will motivate itself to maturity.

Modest yet effective steps emerge in the Near-Term programme. In it, we focus on
activities that are achievable within the next few years. The elements tap into current
capabilities and programmes, seeking to improve international management and co¬
operative structures in preparation for the future.

In the Mid-Term, we focus on more ambitious programmes. Some may require
technology development, but all will have implementation times in the first decade of the
next century. In this Term, we begin fulfilling high priority science objectives, and
envision a continuously operating international solar threat monitoring and earlv
warning system. y

The Far-Term programme is characterised by higher risk, by the use of advanced
technology, and/or by integrated programmes. Elements benefit from and build on the
foundations created earlier. For example, the space threat monitoring and early warning
system begun earlier should be mature enough by this time to create a global forecasting
system, one that provides benefits to developing nations.

As primary areas of scientific interest, we selected the corona, the solar wind, the Sun's
effect on the Earth, and solar theory and model development. In prioritising our
objectives, we found it effective to justify importance based on relevance to the Earth.

In the area of applications, we viewed the Sun as a source of resources and of threats. We
found it useful to search for possible application spin-offs from science missions, for
missions that could be dedicated to a particular application, and for possible future
applications that would require technology development. As our principal focus, we
chose to focus on threat mitigation, by examining ways to improve solar threat
monitoring and early warning systems.

We stress the importance of stereoscopic imaging, of observations at high spatial,
spectral, and temporal resolutions, of long duration measurement to provide information
on physical processes, and of exploring the Sun's polar regions. The corona must be
studied from different observing locations, from closer orbits to the Sun, and by different
means. The Cluster mission must be recovered. The physics of the Sun's interior should
be emphasised more in the Mid- and Far-Terms. Finally, we place emphasis on
monitoring the space weather, forecasting Sun-Earth interactions, and providing early
warning of solar threats. All of these activities should be accompanied by continuing
efforts in theory and modelling.

We found space environment forecasting to be an increasing market. Existing
international solar warning and forecast data distribution networks like the International
Space Environment Service will feed data into forecasts, but the advances needed to
make solar warnings and forecasts relevant to potential users will require capital
investment in hardware, especially in instruments placed between the Earth and the Sun.
Meeting user needs will be essential to commercial opportunities within the larger
government space warning and forecast services.

Improved measurements and models of the space environment will benefit both manned
and unmanned space programmes and thereby constitute a ground for funding. We
envision that humanity will be taking serious steps toward the establishment of manned
lunar outposts or Mars explorations. Study of solar radiation effects on tissue will be
essential A small but important human dosimetry payload flown prior to any such
manned programme is clearly needed. The Ra Strategic Framework has placed such an
investigation in the Mid-Term.

We also suggest that entertainment and education markets can be served by the
conversion of scientific results. We realise that increasing awareness of solar science and
solar-terrestrial interactions beyond the scientific community will foster support for
continued solar exploration and applications. Increasing public interest in the Ra
programme should increase the availability of governmental funding.

There is a trend toward joint ventures between universities and industry. The
universities' research is relevant to industry, and industry funds part of it. We see a trend
where Sun activities are moving from being research driven to product/service driven.

The global political environment within which space activities take place is changing for
a variety of economic, social, and technological reasons. The current international

272 • Ra: The Sun for Science and Humanity

ahon presents both obstacles and opportunities for solar exploration and applications

§e * S 63 7 de< T reasin S national space budgets, and in the relatively low
getary priority currently placed on solar and heliospheric physics and on solar
warning and forecasting services.

Nevertheles 5 , we recognise opportunities for solar exploration and applications. Greater
collaboration leads to multilateral co-operative efforts. Less commercial sectors
experience enhanced co-operation because of mutual payback opportunities and
decreased concern about disproportionate or unilateral technology transfer. The
increasing complexity of the global space infrastructure points to an immediate need for
improved solar warning and forecasting capabilities. Diminishing rivalries between the
various basic and applied sciences facilitate interdisciplinary science missions, and
enhance the possibility of joint science and applications endeavours.

We see the combination of diminishing national space budgets, increased opportunities

”5f°-°P eratlon ' n nd Sr< T?S tech nological capabilities leading to a sustainable
emphasis on smaller, modular, networked spacecraft with prioritised objectives
Disciplinary cohesion, inter-agency co-ordination, international co-operation
applications rationales, and smallsat technology offer us a combination of effective

efforte 53 ° na meanS t0 SUStain 3nd 6Ven inCrease soIar exploration and applications
This situation is ideal for the introduction of Ra.

Once the commitment is made to pursue solar science and applications within this
Strategic Framework, the question of international organising arises. To that end, Ra has
proposed the formation of a Working Group on International Solar Exploration and
Applications (WG ISEA) that synchronises independent efforts in different countries and
helps to combine their output into products useful on a global scale.

We believe that the WG ISEA would be another example of successful organisation, just
as the Inter-Agency Consultative Group for Space Science (IACG) and the International
Mars Exploration Working Group (IMEWG) have been. The WG ISEA, supported bv
small funding from existing sources in the participating countries, has the potential to
unify the scientific community's support for solar science and to facilitate the flow from
science to applications. It could be the forum for bringing into fruition new benefits for
humankind and opening new areas for application and development.

We call attention to the opportune timing with which events will be unfolding during the
next year. ESA will most likely be releasing a Call for Ideas for the M4 mission (part of
the Horizon 2000 Plus programme). The M4 has presently been reserved for a mission
concentrating on the Solar System. The IACG will likely begin the process of choosing its
next focal project. Currently, it has been co-ordinating the International Solar Terrestrial
Physics Program (ISTP). NASA is planning to bring its Sun-Earth Connections Roadmap
to the American space science community for assessment.

Having in place a Strategic Framework dedicated to solar science and applications and a
small but broadly-based international WG ISEA would prove most beneficial to the above
activities. We hope that this report will help to make that happen.

,i

Appendix A

Overview of Sun Related
Missions

!h,H h v S f^ ei i diX “a / evie , we c the past ' current and P lanned missions that are related to

tab^ind -7 and/ ° r tKe Sun *^ arth rela tionship. After a description of the missions, a
table individually summarising each one of them is provided.

A.l Past and Current Missions: Objectives, Characteristics, and
Accomplishments

Interest in the Sun has always existed among the world scientific community but the first
space study of our star begun only in 1949 by the launch of the U.S. NRL V -2 rocket

p tudyin | f ola 5 X ; rays - We must wait U P to 1959 with the launch of Luna 1 by
USSR and of Pioneer 5 by the U.S. to find the first spacecraft carried instruments to study the
bun and its effects. 7

In the following lines we will have a short review of the various probes that have been
launched up to now, what have been their missions and what contribution did they give to
our knowledge of the Sun. The missions are classified in three categories, U.S., Russians or
international missions. International missions consist of the missions planned in co¬
operation between different countries, national spacecraft launched by another country or
national spacecraft carrying non national payloads.

A.1.1 American Missions

Skylab :

The Skylab mission was to prove that humans could live and work in space
periods, and to expand our knowledge of solar astronomy well beyond

for extended
Earth based

observations. Skylab was on an Earth orbit, perigee: 434 km, apogee: 442 km,
inclination: 50.0°.

Skylab was the first U.S. orbiting space station. It was launched on 14 May 1973, from the
NASA Kennedy Space Centre by a Saturn V launch vehicle. Sixty-three seconds after lift-off,
the meteoroid shield designed also to shade Skylab's workshop deployed inadvertently and
was torn from the space station by atmospheric drag. When the meteoroid shield ripped
loose, it disturbed the mounting of workshop solar array number two and caused it o
partially deploy. The exhaust plume of the second stage retro-rockets impacted the partially
deployed solar array and literally blew it into space. Also, a strap of debris from t e
meteoroid shield overlapped solar array number one such that when the programmed
deployment signal occurred, solar array number one was held in a slightly opened position
where it was not able to generate any power.

Approximately 75,000 telescopic images that the Skylab astronauts made of the Sun were
added to the knowledge of our most important celestial body. The images were taken in the
X-ray, ultraviolet, and visible portions of the spectrum. The pictures strengthen the evidence
that the solar corona is more dynamic and complex than previously believed^ On
21 January 1974, for the first time a solar flare had been recorded from beginning to end with
powerful spacebome instruments.

On 11 July 1979, Skylab re-entered the Earth atmosphere. The debris dispersion area
stretched from the south-eastern Indian Ocean across a sparsely populated section o
Western Australia.

Pioneer:

The Pioneer mission consisted of a series of nine spacecraft launched by the U.S. in the sixties
and the seventies to study the Solar System and more particularly to collect scientific data on
interplanetary environment. The information concerning the Pioneer spacecraft is
summarised in table A.l.

Table A.l Characteristics of the Pioneer satellites.

Spacecraft

Launch Date

Type of Orbit

Life-time

Pioneer 5

11/03/1959

solar orbit 0.8 AU

Pioneer 6

16/12/1965

solar orbit 0.8 AU

30 years

Pioneer 7

17/08/1966

solar orbit 1.1 AU

29 years

Pioneer 8

13/12/1967

solar orbit 1.1 AU

26 years

Pioneer 9

08/11/1968

solar orbit 0.8 AU

15 years

Pioneer 10

02/03/1972 |

Interplanetary

Still working

Pioneer 11

05/04/1973

Interplanetary

22 years

Pioneer 12

20/05/1978

Venus orbit

14 years

Pioneer 13

08/08/1978

Venus orbit

Still working

Among all the information gathered by these spacecraft, several were related to the Sun.
Their measurements helped to increase our knowledge about solar wind, cosmic rays,
structure of plasma, magnetic fields, physics of particles and solar flares. The Pioneer probes
were originally designed to last at least 6 months in the space environment, but most of them
have had a life-time of over 20 years.

276 • Ra: The Sun for Science and Humanity

OSO:

J he < ?59 > (0r ^ ng Solar Observatory) mission consist of 8 satellites launched by the U S

^T t , ( ° 75 t0 Earth CirCUlar ° rbit at an altitude around 550 km and inclination around

jJ . 1 he information concerning the satellites is summarised in table A.2.

Table A.2 Characteristics of the OSO spacecraft.

Spacecraft

Launch Date

Type of Orbit

Life-time

OSOl

07/03/1962

Earth orbit 575 km

1 year

OSO 2

Failure

0

OSO 3

08/03/1967

Earth orbit 550 km

2.5 years

OSO 4

18/10/1967

Earth orbit 550 km

4 years

OSO 5

22/01/1969

Earth orbit 555 km

6 .5 years

OSO 6

08/1969

Earth orbit

3.5 years

OSO 7

29/09/1971

Elliptic Earth orbit 330/575 km

3 years

OSO 8

21/06/1975

Earth orbit 550 km

3 years

The OSO mission has collected data on gamma rays. X-rays, solar flares and energy
spectrum. 07

IMP:

( Inter P lanetar Y Monitoring Platform) mission consists of several satellites launched bv
NASA in the seventies. 7

9m P nnn\ aS ^ la “ nched ! n March 1971 on an elliptical Earth orbit with apogee at some
200,000 km, it was designed to study gamma rays (intensity and energy) and to monitor solar
flares. The mission ended on September 1972 due to a failure of the gamma rays instrument.

IMP 7 has been launched in September 1972 to replace IMP 6 and has been carrying on the
same mission up to October 1978. y 6

IMP 8 has been launched by NASA in October 1973 to make measurements on magnetic
fields, plasma and charged particles in the magnetotail, magnetosheath and in the near Earth
solar wind. It was sent to a near circular Earth orbit at a distance of 35 R E . The spacecraft is

still in operation today and provides valuable data very useful to understand long term solar
evolution. b

SOLRAD :

in a , SerieS u°i Satelli , teS launched b y the US. Navy in the seventies to study the Sun.
SOLRAE) 10 launched in July 1971 was posted to an elliptical Earth orbit with apogee at
630 km, perigee at 436 km and inclination of 51°. It was carrying 14 instruments to study
electromagnetic radiation coming from the Sun. SOLRAD 11 A/B were launched together in
Meirch 1976 to a circular Earth orbit at an altitude of 20 R E . They were carrying instruments
to study particles and cosmic rays.

ADDendix A • 777

Voyager:

The Voyager mission was composed of 2 spacecraft. Voyager 1 launched in September 1977
and Voyager 2, launched in August 1977. They were designed to follow on the Pioneer
mission, by studying Jupiter and Saturn and collecting data on the interplanetary medium.
Concerning the Sun, the Voyager spacecraft have carried various instruments tornake
measurements during their journey across the solar system. They have collected data on
particles, cosmic ray! and magnetic fields . By studying their radio emissions, scientist
discovered that the heliopause exists some 90 to 120 AU from the Sun.

Solar Max:

The Solar Maximum Mission spacecraft was launched in February 1980 to a 28° inclined
Earth orbit at an altitude of 500 km. It was designed to provide observations of solar flares
during a period of maximum solar activity and then collected data on solar flares energy,
particle acceleration, CMEs and formation of hot plasma. In January 1981, *ere was a
malfunction and SMM was recovered by the Space Shuttle Challenger in April 1984 and
serviced in orbit. The mission ended in November 1989.

Sampex:

Sampex stands for Solar Anomalous Magnetospheric Particle Explorer, it is the first part of
the U S. SMEX (Small Explorer) programme. It was launched in July 1992 on a 82 inclined
elliptical Earth orbit, at an altitude of 520/670 km. As its mission was to make measurements
on particles, its payload was composed of the most sensitive particles sensors ever flown in
space at that time. Sampex studies the energy, the composition and the charge states of
particles coming from solar flares. It is still in operation today.

Spartan:

Solar Spartan is a mission flown by the shuttle in August 1993. The spacecraft is launched by
the shuttle, deployed in space for a certain amount of time, then recovered and returned back
to Earth for data analysis and maintenance for the next mission. The orbit of Spartan is
elliptical with apogee at 311 km, perigee at 295 km, and an inclination of 57°.

Spartan carries an ultraviolet coronal spectrometer and a white light coronograph to study
solar wind acceleration by examining particles temperature and densities and solar wind

velocities.

A.1.2 Russian Missions

USSR started to study the Sun and the Sun-Earth interactions from the very beginning of its
national space programme. Luna probe series were the first to discover the solar wind. "Luna-
1" on 2 January 1959 was the first lunar flyby. It discovered the solar wind whose existence

was later confirmed by "Luna-3".

In 1960's-1980's satellites from the series "Cosmos", "Electron", "Prognoz", "Intercosmos„
continued the Sun studies. Regular launching of the high-apogee satellites of the Prognoz
series made it possible to conduct unique studies of the structure of the shock wave near the
Earth. The apogee of its orbit is about 200,000 km. "Prognoz-8" studied plasma waves and
accelerated electrons. Intershock experiment carried on the "Prognoz-10" measured the
parameters of the plasma, energetic particles, plasma waves, electric and magnetic fields near
and inside the front of the near-Earth and interplanetary shock waves, The front s structure
and its dependence parameters of the plasma flow in front of the shock wave were also
studied. Oreol satellites launched into polar orbits, made possible to investigate the regions

978 • Ra: The Sun for Science and Humanity

and mechanisms of direct penetration of the solar wind into the magnetosphere The
nfercosmos-Biilgarm-iSOO satellite made research of the physical processes in ^ionosphere

operatinfln 199oT992 f Th ^ the,r , interrelationshi P s - Gamma Space Observatory was
flares [Sagdye^99ll ' re g^tered gamma rays (1011-1012 GHz, up to 29 eV) in solar

CoronaS aimed to study solar activity mechanism, to improve the knowledge about its
internal istructure, to study of magnetosphere-ionosphere processes. "CorfnaS-I" was
“ d H °? * e 02 I" 4 ' ^ Work - iJI “d part of orbit in Z mode": 4

[rthlenlloAm 3 7 COrreSP ° nding t0 high and low solar acdv ity, respectively.

GEOTAIL GT n^TFR??K Ce includes SOHO ' WIND ' POLAR, Interball-1&2,

GEOTAIL, CLUSTER). The primary objective of the mission is detailed study of the energy

Astern ^ and maSS transfer in the critical re 8 ions of the solar wind/magnetosphlre

The "Interball constellation" consists of 2 pairs of satellites (4 altogether)- 2 for

^tpH > ° S p henC f0r aur0ral studies in P olar CUS P S - Each pair consists of a large Russian
l I! 6 Pro 8 n ° z ' M2 and a smaller Czech sub-satellite "Magion". The first pair was
Hunched on 3 August 1995; the second pair was to be launched in August 1996 ^Lisov I.,

A.1.4 International Missions

Helios :

Helios is a German/U.S. mission composed of two satellites. Helios 1 was launched in
December 1974 started working in March 1975 and ended its mission on March 1986 after

Z/Z Cy H e , 11 6 ' time : Hdi ° T S 2 WaS l3Unched in J anuar y 1976, started working on Apri
mbft Z rU V T T S10n -°u J n nUary 198L B ° th ° f them Were P° sted ™ a solar elliptfcal

orbit, m the ecliphc plane with a 0.3 AU (64 Rg) perihelion. F

They were designed to study parhcles, dust, cosmic radiation, magnetic fields, solar wind
mission'to'the Sun. 7 graVitati ° n ’ Up t0 now ' the Helios missi on stills the closest

SIGNE 3:

SrcNE 3 was a small French spacecraft launched in June 1977 by the Russians to an elliptical
Earth orbit with apogee at 519 km, perigee at 459 km and inclination of 50.66° Its has^een
carrying instruments to study gamma rays and solar UV radiation.

ISEE:

1978° totheTl C^rg C epo r lnt Earth ^ 3 ' ^ ^ WaS laUnched ° n Au S ust

was^ompleted ZiZth 0 Solar flares and cosmic gamma rays burst. Its first mission
pleted in 1982 then the satellite was manoeuvred to intercept the comet Giacobini-

Zinner. It flew through its tail in September 1985. In 1990 ISEE 3 was posted to a solar orbit
CMEs^ ^ h ° n ° f 103 AU ' a perihelion of 0.93 AU and an inclination of 0.1° to study

Ulysses :

Ulysses is an international programme, done in co-operation between NASA and ESA. It
was launched in October 1990 towards Jupiter and used its large gravitational field to
accelerate out of the ecliptic plane. Ulysses has made observations of the southern latitudes
of the Sun from June to September 1994, crossed the ecliptic in February 1995 and travelled
trough northern solar latitudes from June to September 1995. As its orbital period is six
years, the next high latitudes observations will be provided in 2001 during a maximum
activity period of the solar cycle.

Ulysses is designed to study and monitor solar flares and detect and localise cosmic gamma
rays bursts. It is the first spacecraft designed to study high solar latitudes.

Yohkoh:

Yohkoh is a Japanese satellite carrying Japanese, American and British experiments. It has
been launched from Kagoshima in August 1991 on an elliptical low Earth orbit with an
apogee of 730 km and a perigee of 570 km. The spacecraft is designed to make observations
of high energy phenomena of the Sun such as flares and others coronal events.

Geotail:

Geotail is a Japanese/American spacecraft part of the International Solar Terrestrial Physics
(ISTP) programme. It has been launched in July 1992 to an elliptical Earth orbit with
maximum apogee at 200 R E , in November 1994 the spacecraft has been manoeuvred to a near
Earth orbit with an 8 R E perigee and a 30 R E apogee. The principal mission of Geotail is to
measure the global energy flow and the transformation in the magnetotail.

Wind:

Wind is an American spacecraft part of the ISTP project. It has been launched in November

1994 to an elliptical orbit with maximum apogee of 250 R E . It will be kept on this orbit for 2
years and then will be moved to the LI Lagrange point. Its objectives are to collect data on
plasma, energetic particles and magnetic fields, to investigate the processes of plasma and to
provide information to be correlated with Ulysses measurements.

SOHO:

SOHO is an ESA/NASA programme part of the ISTP project. It was launched on December

1995 to the LI Lagrange point where it is now able to perform permanent observations of t e
Sun. Its principal objectives are to provide data on the corona, the acceleration of the solar
wind, the solar interior and the solar atmosphere.

Polar:

Polar is the second contribution of NASA to the ISTP project. It was launched in February
1996 to an Earth polar orbit, with a 2 R E perigee and a 9 R E apogee. Its principal objectives are
to measure plasma particles and fields in polar regions, to study auroral plasma and to
provide auroral images.

280 • Ra: The Sun for Science and Humanity

Challenges* 16 *^ MUsi ° nS: ° b i ectives ' Characteristics,

and

SSSSsSi

A.2.1 American Missions

Fast:

Fast is the second SMEX spacecraft, it is planned to be launched in Aueust 1996 hv * P e

ACE:

™ausUW7 to ^ 0mp0sition Exp J? rer > is a NASA mission planned to be launched in

the^galacticmatter^ ' he S ° la * C ° Pr0na - ,he i^net^and
Trace:

u e § i0n And Coronal Explorer) is another part of the SMEX project it is

Timed :

br£fed r r^:s

Lower T^rmosphere/ionosphere) region^^ ^ ° f ** MLTI (Mesos P here and'

HESI:

IMAGE :

2SS " .‘‘a^n « Eatth <>*“ w”™" 5 :

A2.2 International Missions

Cluster:

Cluster was a 4-spacecraft combined ESA/NASA programme part of the ISTP project,
scheduled to be launched by the first flight of Ariane 5. As the flight was not a success, th

four satellites were lost in June 1996.

The Cluster satellites were planned to be posted on an elliptical Earth orbit with an apogee at
19 6 R and aperigee at 4 R, They were supposed to collect data on the Magnetopause, the
polar cusps, the Magnetotail, the plasma sheet boundary layer and the auroral zone.

SAC B:

SAC (Satelites de Aplicaciones Cientificas) is a co-operation mission between NASA and the
TONAF ("Argentina Space Agency). The spacecraft is scheduled to be launched at the end of

orbit will be Earth circular at an altitude of 550 km with an

38° inclination. The satellite will study solar flares, gamma rays burst, X and cosmic rays and
energetic neutral particles.

Plamya:

Plamva mission is the Russian contribution to the U.S./Russia Sun exploration project (see
U S contribution below). It is expected to be launched not later than year 2003 may be m a
combine Proton launch with the Solar Probe spacecraft. Plamya will be posted to an elliptica
Sun orbit with a perigee at 8/10 Rs in the ecliptic plane. It will carry instruments to measure
magnetic fields, particles, cosmic rays and to study corona from a very close point of view.

Solar Probe:

Solar probe is the U.S. proposed mission to be done in co-operation with Russia to investigate
as far P as possible into the Sun corona. Its launch date is expected not later than year 2003 to
an elliptical Sun orbit with a perigee at 4 R s in the ecliptic plane. Its mission is very ambitious
and would help to answer fundamental questions on the heating of the corona and the

creation of solar wind.

Solar B:

Solar B is a co-operation between Japan, USA and the UK to follow on the Yohkoh mission
It is scheduled to be launched in August 2003 on a circular, polar. Sun-synchronous, Earth
orbit at an altitude of 700 km and an inclination of 97.9°. Its objectives will be to provi e co-
ordinated measurements of optical radiation, EUV and X-rays comrng from the Sun to
improve our knowledge of solar activities.

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290 » Ra: The Sun for Science and Humanity _____

Mission Name: GAMMA I Mission Name: GEOTAIL

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296 • Ra: The Sun for Science and Humanity ______

Mission Name: ACE I Mission Name: CLUSTER

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Mission Name: FAST Mission Name: HESI

297 • Ra: The Sun for Science and Humanity Appendix A

298 » Ra: The Sun for Science and Humanity _____

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Ra: The Sun for Science and Humanity Appendix A

Appendix B

SAUNA Mission Data

The following information on the SAUNA mission is included here:

• Reference low-thrust transfer trajectory.

• Detailed budgets (mass, power, cost) supporting Table 9.6.

• Communication link budget analysis.

B.l Reference Low-Thrust Transfer Trajectory

These input parameters were fed to SKYNAV [Appendix C]:

Positioning Manoeuvre/Initial Mass 300kg/Exhaust Velocity 45911 m/s/Thrust
0.200N/Solar-Electric Propulsion Thrust Radius l.OOrr/Thrust Exponent
1.700/Initial Delta-V O.OOm/s/Initial orbit pericenter radius 0.722rr/excentricity
0.16/Final orbit pericenter radius 0.200rr/Excentricity 0.000/Angles [deg]: Initial
Orbit pericenter/Node Axis: 180.00, Node Axis/Final position 1970.00 fall other
angles 0.000]

The results of the SKYNAV optimisation were:

• Total Delta-V Requirement: 34.19 km/sec

• Total travel time: 507 days, whereof first 90 days spent coasting

Figure B.l shows a graphical output as provided by SKYNAV. The total delta-V is higher
than for impulsive (Hohmann) manoeuvres since more losses are incurred as a result of
continuous thrusting against various external influences. The transfer time is a direct
function of the number of spiralling turns needed to reach the 0.2 AU orbit (in this case 5.5
rounds).

Figure B.l SAUNA Reference Trajectory Characteristics

304 • Ra: The Sun for Science and Humanity

B.2 Detailed SAUNA Budgets

SAUNA BUDGET
Item

Al. Primary Structure
A2. Secondary Structure

A. Structure

Power Cost

(US$ M) Remarks

Page 1/2
Coefficients

> 0.00| _2.00i_

!_0 00 2.00 for instrument support, etc.

10.00 4.00

Bl. Main Ion Thruste r
B2. Propellant Tank

B. Propulsion System

20.00 6300.00' _ 2.00 Ion p ropuls ion UK-25E at 0.2N j

10 00_0-00i 3 00 Sizing needed for specific propellant

1 [6300.00 5.00

Cl. Solar Panels

C2. Batter ies_

C3. Power Control Electronics

C. Power System

31.61

: 0.00

3.00

0.00

1.00

0.20

2.00 M=0.016Ptot/4; ISS/Lockheed Martin Flexible

0.50^ Storage capacity TBD

0.25

Dl. Star Trackers (2x)

0.58

12.00

D2. Inerhal Measurement Unit

0.20

24.60

D3. Ion RCS Theusters (6x)

12.00

0 00

D4. Reaction Wheels

14.10

35.00

2.00 STSC [L LNL 19961

1.0 0 _ _ _ ;

1.80 A ssume RCSOR Main, never both
2.00 [Ithaco 1996] ]

D. Attitude & Orbit Control

E. Thermal Protection

FI. High Gain Antenna _

F2. Low Gain Antenna
F3. Antenna Drive mechanism
F4. X-band Transponders 2x
F5. S-band Transponders 4x

15.00

0.00

5 00 2m dish [Appendix F]

4.00

0.00

2.00 4x S-band [Appendix F]

2.00

0.00

2.00

2.00

110.00

2.00 Pluto

4.00

640.00

2.00

Gl. CPU

1.20

G2. Mass Memory

0.60

G3. Data Bus

0.20

750.00 113.00

G. On-Board Computer

12.00 [9.00 6.00

H. Subtotal Spacecraft Bus:

1188.49 17165.80 48.30 [A+B+OD+E+F+G1

AoDendix B • 3ns

Cost

SAUNA BUDGET
Item

(US$ M) Remarks

[ Page 2/2
Coefficients

11. Plasma Analyzer

3.70

2.70 3.00

12. Magnetometer

2.50

1.50 3.00

13. EUV Telescope

6.00

5.00 3.00

14. X-Ray detector

2.00

4.00 3.00

15. Dust detector

0.50

2.00 0.50

16. Common electronics box

3.00

i 1

; 3.oo o.5o! i

Kl. Harness mass _ 20,62 ____

K2, Harness power loss _ 35.92 _

K3. Harness cost _0

LI. Subtotal Dry Mass 1226.81 17219.92 61.34

L2. Prope llant (52.5% of wet masi 250.68 n.a. 0.63

M. Subtotal Wet Mass _ 477.49 7184.00 61.97

N. Margin _ 47.75 718.40 12.39

O. Total Wet Mass Spacecraft: 52 5.24 .n.a. 74. 36

P. Launcher adaptor; 15.76 n.a. 0 50

Q. LAUNCHED SUBTOTAL: 1541.00 17902.40 74.86

_ Percentage cable mass of above: _

_ 18V Bus, Average Power Loss %: _

0.04 Cost harness mass per kg, $_

i : n±g__

52/48 fuel/mass ; Xe gas $/kg: _

7 } [LI+L2] ____L_

9 i 10% for mass & power, 20% for cost

- | - -—— ---

6 [M+N]

- 1 --

_13% of wet mass_ i

6 [O+Pl I

Annual MODA 3.8M*5yrs (MOCM

T. SAUNA Predevelopment (SPP)

U. Subtotal Cost:

Q+R+S+T

TOTAL MASS, POWER, COST:

541.00

7902.40 194.00

IU+V]

1

Launched Mass/C3 Function for Delta II (7925)

Current C3 Pointer:

Mass (kg)

C3 (km2/s2)

277.1

75

308.3

70

379.7

60

465.7

50

XXXXXXXXXXXXXXX

569.9

40

697

1 30 !

B.3 Link Budget Analysis for SAUNA HGA

DIRECT RF LINK (HGA)

Concept

Value

Frequency (GHz)

8.40

Transmitter RF power (W)

40.00

Transmitted power (dBW)

16.02

Transmitter losses (dB)

-1.50

Transmit antenna diam (m)

2.00

Transmit antenna gain

17023

Transmit antenna gain (dB)

42.31

Tx 3 dB Beamwidth (deg)

1.25

EIRP (dB)

56.83

Sun apparent angle (deg)

0.27

Distance from Sun center (AU)

0.20

Max Margin Angle (deg)

11.04

Margin angle (deg)

11.00

Min SEP Angle (deg)

11.27

Path length

1.02

Path loss (dB)

-274.59

Implementation loss (dB)

-2.00

Polarization mismatch (dB)

-0.01

Receive antenna diam (m)

30.00

Receive antenna gain (dB)

65.83

Rx 3dB Beamwidth (deg)

0.0833

Antenna pointing loss (dB)

-0.35

Antenna noise temperature (K)

50.00

Feeder noise temperature (K)

270.00

Connection loss (dB)

-1.00

Receiver noise temperature (K)

50.00

System noise temperature (K)

145.25

Noise spectral density (dBW/Hz)

-206.98

G/T (dB/K)

44.21

C/No (dBHz)

52.70

Max Info. Data rate (kbps)

16.00

Channel Code rate

0.44

Overall Data Rate (kbps)

36.59

Eb/No (dB)

7.06

Implementation loss (dB)

-0.50

Required Eb/No (dB)

2.50

Margin (dB)

4.06

Appendix B • 307

Link Budget Analysis for SAUNA LG A

DIRECT RF LINK (LGA)

Concept _

Frequency (GHz)

Transmitter RF power (W)

Transmitted power (dBW) _

Transmitter losses (dB)

Transmit antenna diam (m)

Transmit antenna gain (dB)

Tx 3 dB Beamwidth (deg)

EIRP (dB) _

Path length

Path loss (dB) _

Implementation loss (dB)

Polarization mismatch (dB)

Receive antenna diam (m)

Receive antenna gain (dB)

Rx 3dB Beamwidth (deg)

Antenna pointing loss (dB)

Antenna noise temperature (K)

Feeder noise temperature (K)

Connection loss (dB)

Receiver noise temperature (K)

System noise temperature (K)

Noise spectral density (dBW/Hz)

G/T (dB/K) _

C/No (dBHz) _

Max Info. Data rate (kbps)

Channel Code rate

Overall Data Rate (kbps)

Eb/No (dB) _

Implementation loss (dB)

Required Eb/No (dB) _

Margin (dB)

Value I Observations

2.20 S-band

160.00 SS: 640 W DC 12 K

22.04

1.20

-264.40

270.00

- 1.00

50.00

145.25

-206.98

32.57

0.441 RS(255,223)+Conv.(7,.5)

3.43

7.08

-0.50

2.50 RS/Viterbi BER+le-6

4.081 Conv. BER=le-6

308 • Ra: The Sun for Science and Humanity

Appendix C

C.l In-situ Instrumentation for Various Missions

Table C.1.1 Minimum Solar Mission Instrument and Measurement Parameters

Measurements

Spectral

Spectral

Integra tio

Observatio

Mass

Power

Cost [$m]

Paramete

Resolution

n Time

nal

[kg]

[watts]

r

cadence

Plasma

Plasma Electron

Energetic

Particles

.01-30

keV

30% 1

10 s

100 s

4

.01-30

keV/Q

7.05%

Ions

>100keV

Electrons

25 keV -
4.5 MeV

Protons

400 keV -
45 MeV

Alpha

1.3 -180
MeV

Wave

1Hz-

10kHz

+ /-0.2G

[W.I.Axford, 1996, "Minimum Requirements For A Solar Probe Mission", Adv.Space Res Vol

17, No. 3, pp. (3)85 - (3)90]

Table C.1.2 Russian FIRE Spacecraft Payload

Experiment

Mass [kg]

Power

[W1

Bit rate
[kbit/s]

Requirements

Plasma Analysis

6

6

2x2?t FOW ions

Electron sensor on the boom

Sun-directed hole required

Aspect precision 1°,
knowledge 0.2°

Magnetometer

4

3

0.5-5.0

2 sensors on the boom

Energetic Particles

3.5

4.5

0.3-1.0

Measurements in 4 directions

Plasma Waves

mm

6

<15

2 booms lm each

Current collectors on thermal screen

Interplanetary

neutrals

■

2

Neutrons and y

3.5

3.5

16/day

Table G1.3 German M3 Mission

Instrument

Mass Power

[kg] [W]

Telemetry

[bits/s]

Solar Wind Plasma
Particle Analyser

Plasma Wave

Electric Field Instrument

Digital Wave Processin

Suprathermal Sensor

.4030

2500

Solar Energetic Particle
Analyser

3.5

4

Dust

Detector for
Interplanetary Dust
Particles

3-D Ion Velocity
Spectrometer

Proton Alpha Sensor

Thomson Parabola
Analyser

Electronics box &
connectors

Tilt table <5c electronics

Total

Ions

Heavy Ion Analyser

3-D Elect Velocity
Spectrometer

Magnetometer

Time

Resolutio

Field Strength

n

[S]

O.lnT-O.lG

20 - 1000 keV

50keV-50MeV

electrons

4keV-10MeV

3.5

100

3.0

4000

5200m

lmT/32nT

6400nT/2nT

3200nT/0.1nT

256nT/8pT

10k-

50M

1 - 500 &
500-50M

• Pa- Tko Qnn for Siripnrp and Humanitv

[V.N.Oraevsky and V.D. Kuznetskov, 1996, "Instrument For The Russian Solar Probe Mission", Adv.Space Res Vol 17 No 3 pp

(3)103-(3)1101 ' ’

Table C.1.4 YOHKOH

Instruments

Resolution

Time Resolution

1 Spectrometer j

Bragg Crystal Spectrometer

(BCS)

Sxv (5.0160-5.1141A)

3.232 mA

0.125

Ca xix(3.1769A)

0.918 mA

Fexxv(1.8509A)

0.710 mA

Fexxvi(1.7780A)

0.565 mA

[Y.Uchida, The YOHKOH Mission, Solar Physics , Vol. 136, No. 1,1991, pp. 69 - 88]

Table C.1.5 Russian Solar Probe Mission Complex of Electromagnetic Remote and Itt-situ
Measurements (CERIM)

Instruments

Mass [kg]

Power [W]

Data [kb/sec]

■

Magnetometer

4

3

4

Electron Gun

3

5

5

Solar Radiospecttometer

2

5

2

100k -30M

SOHO

CDS (Coronal Diagnostic Spectrometer)

CELIAS (Charge, Element, and Isotope Analysis System)

COSTEP (Comprehensive Suprathermal and Energetic Particle Analyser) from the University of Kiel, Germany
ERNE (Energetic and Relativistic Nuclei and Electron experiment)

GOLF (Global Oscillations at Low Frequencies) from the Institut d'Astrophysique Spatiale, France
SWAN (Solar Wind Anisotropies) from FMI, Finland
[E sa Bulletin , ESA/ESTEC, pp. 96-105, 1996]

Solar Maximum Mission

Coronagraph/Polarimeter 4465 - 6583A, 1.5-6 sq.solar radii fov. 6.4 arcsec res.

Ganuna-ray Spectrometer Nal(Tl), 0.01-100 MeV in 476 channels, 16.4 s per spectrum

Active Cavity Radiometer Irradiance Monitor 0.001 -1000 micrometer solar flux

C.2 The Solar Sailing Trajectory Program

C.2.1 Introduction

During the ISU summer session, a software was created to compute solar sail trajectories for
spacecraft already orbiting in a circular fashion around the sun. The code of the program was
written in FORTRAN and was based on previous studies made during the ISU Summer
Session Program of 1994 in Barcelona. Berry Sanders helped in acting as the scientific adviser
for the Sailing program.

You will find, in the first section of this appendix, the complete code in FORTRAN, and the
main output file that was generated with the program for the following values: mass=250kg,
sail area=9000 m2, starting distance from the sun: 150e6 km, and an angle of attack of 45
degrees for the sails, relative to the incoming solar pressure.

A nnortrl iv m 1

C.2.2 FORTRAN Code

cccccccccccccccccccccccccccccccccccccccccccccccc
C Solar sailing ISU-96 Solar Probe Design Project
C SAILING.FOR

C Design & Programming: Marc Abela
C Creation: August 25 1996
C Completion: August 27 1996

C Comments, suggestions and scientific advices:

C Berry Sanders

CCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCCC

Implicit double precision (a-h,m,p-z)

Dimension y{4)
logical kop
external esail

common mass,sailarea,alpha,muesun,asail,ar,al

C Main constants used through out the program

n«4

pi-3.1415926535
muesun*1.327e20

write(6,*) * Enter the mass of spacecraft (kg)’
read(5 ,*) mass

write(6,M 'Enter the solar sail area (m A 2)’
read(5,*) sailarea

write(6,*) 'Enter the starting distance (km) from the Sun*
read{5, *) distance

C Convert distance from km to meters
distance-distance*1000

write(6,*) 'Enter the angle of attack of the sail (degrees)’
readl5,*) alpha

C Convert alpha from degrees into radians
alpha*alpha*pi/180

write(6,M 'Enter the complete duration of the trajectory (days)'
read(5,*) tinte

C Convert tinte from days into seconds for the duration of the flight
tinte * tinte*24*60*60

write(6,*) 'Computing. Please wait... 1

C Open the file pipes and initialize the content

open(unit-7,file='sailingl.out',status*‘old’)
open(unit*8,file* 1 sailing2.out',status* 1 old 1 )
open(unit-9, file*'sailing3.out',status*'old')
write(7,*) 'time',' r' ( lambda',' vt', ' vr'
write(8,*) 'time 1 ,' r',' asail',' ar 1 , 1 al’

write(9,*) 'time',' rp',' ra'

C Kop is for the integration routine no stdio (Kop means head in Dutch)
kop ■ .false.

C Tintl is the beginning of the integration step
tintl * O.OdO

C Tint2 is the end of the integration step
tint2 * O.OdO

C Step size for the integration
3tepnum * 5000.OdO
time « 0.OdO

C Initialization

C y(l) is set to the distance from the sun (in meters)

C y(2) is set to the radial velocity (in meters per second)

C y(3) is set to the starting angle value (in radians)

C y(4) is set to the angular velocity (Vt/R) (in radians per second)

y(1)-distance
y(2)=0.
y(3) *0 .

y(4)*sqrt{muesun/distance* * 3)

C While end of integration is not reached
111 if (tint2,It.tinte) then

tint2*tintl*stepnum
time*tint2

C The next line calls in the Runge-Kutta integration routine
call ruks(n,tintl,y,tint2,1,0,esail, kop)
if (i .ge. 150) then
i*0.OdO

vi-sqrt{(y(4)*y(1))**2 + (y(2))**2)

h0-y(l)*vi**2./muesun

a-y(l)/(2.-hO)

gai-atan(y(2)/(y(l)*y(4)))

enc-sqrt(1-hO*(2-hO)Mcos(gai))**2)

hap*a*(1+enc)

hep-a*(1-enc)

C Time is converted in days, distance in AUs, velocity in km/s, alpha in degrees
timeindays*time/(24*60*60)
write(7,*) timeindays,y(1)/149.6E9,

$ y(3)*180/pi,y(4)*y(l)/1000,y(2)/1000

write(8,*) timeindays,y(l),asail, ar, al
write(9,*) timeindays, hep/149.6e9, hap/149.6e9
endi f
i « i+1
tintl=tint2
goto 111
endif

C Close the file pipes
Close(unit*7)

Close(unit*8)

Close(unit-9)
stop
end

Subroutine esail(time,y,f)

Implicit double precision (a-z)
dimension y(4),f<4)

common mass,sailarea,alpha,muesun,asail,ar,al

C This routine is called by the integration routine (Runge-Kutta)

C Note that muesun is equal to 1.327E20 (gravity constant of the sun)

C y(1) is R

312 • Ra: The Sun for Science and Humanity

C y(2) is Rdot
C y (3) is Lambda
C y(4) is Lambdadot

C ar is acceleration in radial direction away from the sun
C al is acceleration in tangential direction
C asail is the acceleration provided by the solar sails

C These are the default acceleration values for attraction from sun body
C ar=-(muesun)/(y(1)**2.)

C al=0

C Note that yl in the equation needs to be expressed in AU 1 s

C Please refer to the following book for more information on Solar Sails and
C mathematical equations

C Space Sailing, Jerome L.Wright, Gordon and Breach Science Publishers

C 1994

C T is computed as the thrust generated due to the presence of the sail
T*0.9216E-5M (Cos (alpha} )**2 . ) *sailarea/ ( (y(l) /149 . 6e9) **2 . )

C asail represents the acceleration generated by the above thrust
asailaT/mass

ar=-(muesun)/(y(l)**2.)+asail*Sin(alpha)
al*-asail*Cos(alpha)

f (1) *y {2)

f(2)*ar+y<1)*y(4)**2.
f(3)*y(4)

f(4>=al/y(l)-2.*y{2)*y(4)/y{l)

return

end

C.2.3 References:

ISU Summer Session '94, Solar System Exploration Design Project, ISU, Summer 1994

C.3 The MIDAS Trajectory Optimisation Program

C.3.1 Introduction

During the ISU summer session, the MIDAS software was introduced to the Solar System
Exploration design project to assist in the computation of interplanetary trajectories. It was
therefore installed on a Sun workstation and several students were trained to use it by Stacy
Weinstein from JPL, USA. MIDAS is a program developed at JPL. It is capable of optimising
interplanetary trajectories by using a patched conic approximation [Sauer, 1994], It is written
in FORTRAN and runs on both DOS and UNIX computers. MIDAS can compute direct
flights, one or more gravity assists and deep space manoeuvres to a selected target. It is
possible to perform a fly-by, a rendezvous or an orbit around the target.

C.3.2 The Patched Conic Approximation

MIDAS uses a patched conic approximation for the computation of interplanetary
trajectories. The patched conic method uses ideal Kepler orbits for the different phases of the
flight. MIDAS then connects them at the beginning and end points.

To illustrate this, let us take the example of a flight to Mars. The departure from Earth is a
hyperbolic Kepler orbit around the centre of the Earth. The flight from Earth to Mars is an
ellipse around the Sun, while the arrival at Mars is again a hyperbolic orbit around the centre
of Mars. In this case MIDAS computes the different Kepler orbits roughly. Midas then
changes the orbits to connect them at the Earth and Mars orbits to form one consistent
trajectory. Of course, the patched conic approach is not limited to two bodies and three
trajectory parts, more can be included in MIDAS to form gravity assist trajectories and
multiple flybys. MIDAS also has the possibility to include deep space manoeuvres to make
gravity assist trajectories possible.

C.3.3 Possibilities of MIDAS

MIDAS can compute trajectories with up to eight deep space manoeuvres and several gravity
assists from larger bodies. It is also possible to visit one body more than once. Asteroids and
other small bodies are included in a separate table which can be called upon by MIDAS so
flights to nearly all the small bodies can be computed.

The input to MIDAS is a file composed of several lines, describing the starting body, the
target, the intermediate bodies which will be visited and the time frame in which the flight

Appendix C • 313

will occur. MIDAS will find the optimal launch date around the dates given by the user and
it will vary the flight time within the limits given by the user, that way, the user can pick the
optimal flight time. Also, the user has to specify whether he wants to make a fly-by at the
target or orbit it. In the last case he also has to specify the orbit.

The output of MIDAS can be given in different forms, both abridged and extended output
files can be generated. With the help of the K-plot program, it is also possible to generate a
plot of the trajectory.

A separate program called LV can be used to assess the interplanetary performance of
different Western launch vehicles with and without upper stages.

C.3.4 Optimisation Method Used in MIDAS

MIDAS uses a gradient search method to optimise the trajectory. It minimises the total
velocity of a trajectory by taking an initial estimate and computing the gradient towards the
lowest velocity for the mission. It then starts searching in the direction of this minimum for a
given duration of the flight. The danger of this method is that the program can trace a local
minimum and that there might be another global minimum with a lower velocity
requirement. Therefore, the result have to be interpreted with some care.

C.3.5 Concluding Remarks

MIDAS was used quite extensively in the Ra design project for the different feasibility
studies, using Venus, Mercury, and Jupiter flybys and proved to be a very valuable tool for
our project. However, due to the fact that it is an expert tool, the output was often difficult to
interpret for a person with little experience with the program.

C.4 The SKYNAV Trajectory Optimisation Program

C.4.1 Introduction

During the ISU summer session, the SKYNAV software was introduced to the Ra design
project to assist in the computation of interplanetary trajectories. It was installed on a laptop
computer and several students were trained to use it by Berry Sanders from Bradford
Engineering. SKYNAV is a program developed by Ingenieurbuero "Dr. Schlingloff", from a
program originally developed for European ion propulsion missions. It is capable of
optimising interplanetary low thrust trajectories. It is written in FORTRAN and C and runs
on DOS computers. SKYNAV can compute low thrust flights and deep space manoeuvres to
a selected orbit.

C.4.2 Possibilities of SKYNAV

SKYNAV can compute low thrust trajectories with varying thrust and specific impulse
throughout out the trajectory or an orbiting spacecraft using electric propulsion (EP) as its
main source of thrust. It is also possible to visit more than one planetary body (for example,
when rendezvous with asteroids and other small bodies or the solar system).

The input to SKYNAV is a file composed of several lines, describing the nature of the initial
and final orbits, the mass and the propulsion of the spacecraft thought out its trajectory.
SKYNAV will find the optimum low thrust trajectory between the given initial and final
orbits based on the Hamilton-Lagrange optimisation theory. The user can therefore pick the
optimal flight time and propulsion management schemes. The user can also specify whether
he wants to make for example a plane change or modify other orbit parameters.

314 • Ra: The Sun for Science and Humanity

The output of SKYNAV can directly be given in graphical and numerical forms. It is
therefore possible to generate databases as well as graphs of the trajectories selected by the
software. Due to the direct graphical output, it is feasible to visualise and assess the nature of
the optimum trajectory for instantaneous feed-back.

C.4.3 Optimisation Method Used in SKYNAV

SKYNAV uses the Hamilton-Lagrange theory to optimise trajectories. It minimises the total
Delta V of the selected trajectory by taking an initial estimate for the Lagrange multipliers
and by computing simultaneously the equations of motions which describe the trajectory and
the optimality conditions. It then starts searching in the direction of a minimum for a given
travelled angle of flight.

C.4.4 Operation of SKYNAV

The user needs to first set up the nature of the problem and then initialises it. The program
will now solve the problem using impulsive shots. The user will then, in successive runs, try
to lower the thrust in order to create a low thrust trajectory. This operation is a iterative
process that needs to be carried all the way until the desired thrust level is achieved. When
the given mission angle is not sufficient to achieve the required Av, the mission angle should
be enlarged accordingly.

C.4.5 Concluding Remarks

SKYNAV is a very useful program for the computation of optimal low-thrust trajectories. We
used SKYNAV to establish (among other things) the feasibility of the SAUNA mission.
Although other advanced features like gravity assists are not included among the options in
this program, its simplicity of use makes it ideal for simple feasibility assessment of low-
thrust interplanetary missions.

C.5 Data used for trajectories

Table C.5.1 Data of Potential Gravity Assist Planetary Bodies

Planet

Equatorial
radius (km)

I X IT

Sidereal

period

Escape velocity
km/s

Mass

(Earth = 1)

Mercury

2335

57.9

57.9 days

4.2

0.06

Venus

6200

108

224.7 days

10

0.82

Earth

6385

149.6

365.3 days

11

1

Mars

3880

227.7

687.0 days

6.4

0.11

mKBSSSWM

71500

777.8

11.86 yr

59.7

318

Table C 5.2 Common Distance Units

Acronym

Name of the Unit

Value

Definition

AU

Astronomical Unit

149,689,534 km (approx.:

150 M km)

Mean distance between the
Earth and the Sun

Rs

Solar Radius

696,000 km (approx.: 0.7 M
km)

Radius of the Sun

Appendix C • 315

C.5.1 Plamya

This Russian mission is being proposed to fly with the US's Solar Probe on a joint mission
called Fire. Plamya and the Solar Probe would be launched on a Proton launcher for a
Jupiter gravity assist. The Plamya trajectory is very similar to the solar probe's
(perpendicular to the ecliptic), except perihelion will be ~10 solar radii (-0.05 A.U.)

C.5.2 SOHO

The Solar and Heliospheric Observatory ( SOHO) is designed to study the internal structure
of the Sun, its extensive outer atmosphere, and the origin of the solar wind - the stream of
highly ionised gas that blows continuously outward through the Solar System. It will also
study the vibration of the Sun with a very high spatial and temporal resolution. SOHO was
launched on December 2, 1995, by an Atlas 2AS/Centaur from Cape Canaveral, Florida,
USA. The spacecraft has been injected into a Halo orbit around the LI Libration point of the
Earth-Sun system, approximately 1.5 million km sunward from the Earth., requiring a Av=
1.35 km/s. Halo orbits around the LI point are unstable, and small correction manoeuvres
must be applied to prevent excessive departure from the nominal orbit. SOHO's orbital
period is six months.

C.5.3 Solar Probe

The Solar Probe mission, studied by NASA, is a planned fast flyby of the Sun. The trajectory
relies on a Jupiter gravity assist to crank the orbit perpendicular to the ecliptic and approach
within four solar radii (-0.02 A.U.). Due to the gravity assist all the way out at Jupiter, the
time from launch to perihelion is over 4.5 years. However, the primary advantage is that the
spacecraft propellant required is very small, only for orbit corrections and attitude control.
The velocity increment required for injection to Jupiter is - 8.8 km/s which is provided by
the launch vehicle and upper stage.

C.5.4 Ulysses

The Ulysses Mission is the first spacecraft to explore interplanetary space at high solar
latitudes. Its primary mission is to characterise the heliosphere as a function of solar latitude,
with particular emphasis on the regions above the solar poles. The spacecraft was launched
on October 6,1990, by the Shuttle Discovery with two upper stages, during the 5-23 October
Jupiter window.

Since direct injection into a solar polar orbit from the Earth is not feasible with chemical
propulsion, a gravity-assist is required to achieve a high-inclination orbit. For that reason,
Ulysses was launched at high speed (Av=11.4 km/s) towards Jupiter, after being deployed
from Discovery in a 300 km circular low-Earth orbit. Following the fly-by of Jupiter in
February 1992 and the resulting large gravity assist, the spacecraft was injected into an orbit
out of the ecliptic plane with a perihelion of 1.34 AU. The spacecraft is now travelling
northwards in an elliptical heliocentric orbit inclined at 80.2 degrees to the solar equator.
Ulysses achieved its maximum southern latitude of 80.2 degrees on September 13, 1994. It
travelled through high northern latitudes during June through September of 1995. Ulysses'
orbital period is six years.

C.6 Optical Communications

C.6.1 Concept

Free space optical communications offer a substantial increase in link capacity. This increase
comes as a result of the much smaller wavelength associated with optical carriers. Smaller

316 • Ra: The Sun for Science and Humanity

wavelengths result in narrower transmitted beam divergences and hence more concentration
of the power on the intended target receiver. For example, if one compares X-Band (3 cm) RF
with visible light (0.5 micro-m), the ratio of power concentration is 95 dB. Not all of this gain
results in link advantage however. The basic process of photo detection is less efficient than
for RF due to quantum noise effects. When these effects are taken into account, the resulting
overall link benefit is typically about 71 dB [Shaik 1995] over an RF system.

C.6.2 Environment interference

The Sun is, of course, a very bright source from the Earth's point of view in the optical and
near-infrared bands. Direct Sun is a powerful source of optical radiation. The spectral
irradiance of the Sun peaks at 460 nm, as expected from black body considerations, and
decreases with increasing wavelength [Shaik 1995]. In the far infrared, the Sun output
decreases to an acceptable level [Mendell 1996] It is thus possible to use a laser source (e.g.
C0 2 ,10.6 micron) that will possibly have a lower background noise.

C.6.3 Implementation

Because of the narrow beamwidths of optical systems, and due to the finite speed of light,
optical communications signals must be pointed ahead of the apparent location of the
intended target receiver so that the transmitted beam will intercept it. The magnitude of this
point-ahead angle depends on the relative cross velocity of the two communication terminals
and can be as large as 500 micro-radians in some planetary applications [Lesh 1992], The
transmission path steering mirrors are used to introduce this offset angle.

It is highly desirable to develop optical communication systems around Fraunhofer lines
where the Sun s spectral irradiance is substantially low. Ffowever, assuming present
technology, it is difficult to see how this information can be used to an advantage. It is a
technical challenge to produce an optimal match between the laser wavelength and a strong
Fraunhofer line, and in addition, most practical lasers for optical transmitters have broader
line widths than the fine atomic dark lines in the Sun's spectrum [Shaik 1995],

By far the most complicated problem is to have a coherent beam. Firstly, the frequency
accuracy of the diodes is of paramount importance. Presently, the stability of the frequency
output of most diodes is questionable. Furthermore in a diode array, the coherence of the
signal will be tricky as the exact timing of signal generation in each diode will be hard to
implement.

Coherent detection will be equally difficult. Any drift at the emitting end will have to be
closely calculated. The Doppler effect of a probe moving at high velocities throughout the
solar system while the Earth is orbiting will induce a trajectory-dependant Doppler shift
which will have to be known precisely in order to have (1) adaptive filtering and (2) adaptive
coherent signal recombination.

C.6.4 Acquisition

Contrary to a typical near-Earth crosslink the one-way beam propagation time for deep
space communications can be from several tens of minutes to several hours. Furthermore, by
the time the beam reaches the other terminal, the original platform may no longer be in that
location. Thus, an acquisition, tracking and pointing strategy which does not require two-
way beam propagation is needed [Lesh 1992]. Fortunately, at planetary distances, nature can
provide the necessary spatial references in the form of the solar-illuminated planets
themselves. As soon as there is a clear path to the Earth, a telescope can be pointed in that
direction. Once Earth is acquired, a two-dimensional detector array can resolve the image
and, by locking onto that image with sufficient accuracy, adequate knowledge is available to

point and fire the return communication laser beam at the Earth receiver. Since the solar-
illuminated Earth image is always available at the distant location (except when going
behind the Sun) the entire acquisition tracking and return beam pointing process can be
accomplished in a relatively short time period (likely under 30 seconds)[Lesh 1992]. This
permits almost instant communications as soon as line of sight is established.

C.6.5 Ground segment

There are no ground segment yet available for optical communications, or more precisely for
deep space communications. There are two possibilities: either the receptors are placed in
space, or on Earth. Spacebome receptors have already been discussed in Chapter 6.

To correct for atmospheric effects for Earth based receptors, one would have to use a variety
of techniques. The sitting of the receivers on high ground would in the first place alleviate
for atmospheric absorption. Observatories with clear and dry skies have been in place for
quite a number of years. Filtering of the incoming signal is necessary. Again, astronomy has
been filtering light for years, and adaptive filters to work at the desired frequency pose no
problems. Adaptive optics will be required for atmospheric turbulence.

A ground based interferometry system could be developed to increase the amount of energy
collected. Early designs [Lesh 1994] for interstellar missions had 10-m diameter telescopes at
the receiving end. Reducing this by one order of magnitude, it is possible to imagine a global
receiving network of multiple 1-m sized telescope. A number of solar observation telescopes
are in place around the world today. A number of old astronomical telescopes are still quite
usable. They could be converted for our use with proper filtering, conducting observations
of the sky by night and communications by day.

If interferometry is used on the receiving end, the position of the receiving antenna becomes
critical, within fractions of the wavelength. In our case, optical wavelengths are in the order
of micrometers. This means that any variation in the position of the telescopes (solar photon
pressure for orbiting receivers, any type of ground movement around the Earth-based
telescopes) would be sufficient to lose signal lock. Of course, a variety of techniques could be
used to counter movement, especially on the ground. Low-temperature physicists have long
ago recognised that problem and this type of movement accuracy is achievable.

Time tagging of data in interferometry is also of paramount importance. It remains to be
seen (calculated) if the best atomic clocks/GPS systems available now can provide the
required timing.

318 • Ra: The Sun for Science and Humanity

C.7 Ground Station Block Diagram

Fig. C.7.1 Typical ground station block diagram

The German Space Operation Centre, that is operated by the DLR, offers a good example of a
typical earth station that is an alternative to the DSN for supporting the Ra missions. The
centre located at Weilheim/Lichtenau offers a variety of antennas and associated buildings
that can provide support for missions using S-band for the two way links and X- or C-band
for the downlink. It consists of a 30 m antenna and two others of 15 m. [DLR Ground
stations at WEILHEIM, Deutsche Forschungsanstalt fur Luft- und Raumfahrt e.V., WWW]

Appendix C • 319

Appendix D

Costing

Cost database of different Near-, Mid- and Far-Term missions are presented in this
appendix, in the following order :

• Cost per mass and Total Payload mass vs location of mission (Cost
missions).

• Cost of Launchers (CostLaunchers).

• Cost matrix of CLEMENTINE 2 (CostCLEMENTINE2).

• Cost matrix of CLUSTER (CostCLUSTER).

• Cost matrix of FAST (CostFAST).

• Cost matrix of SOHO (CostSOHO).

• Cost matrix of TIMED (CostTIMED).

• Cost matrix of FIRE Mission, Solar probe (CostFIRE Solarprobe).

• Cost matrix of FIRE Mission, Plamya (CostFIREplamya).

• Cost matrix of FIRE Mission, Plamya and Solar probe (CostFIRE
Plamya&solarprobe).

• Cost matrix of RA Application (CostRA-Application).

• Cost matrix of SAUNA (CostSAUNA).

The * represents missing values.

FIRE

Cost per mass and Total Payload mass vs location of mission

k$cost per mass and Total Payload Mass vs Location of mis

5 Kg3000

a

E

•o Kg2500

5

>»

2. Kg2000

a

® Kgl500
Kg 1000
Kg500
KgO

Reference: Isakowitz.S.J, International Reference guide to launch Systems, AIAA 1995.

Cost matrix of CLEMENTINE 2

|HB»tiihhi:iwirowa

MISSION COST Break down Sturcture

□ GROUND SEGMENT
■ LAUNCHER

□ SPACE SEGMENT

SPACE SEGMENT COST Break down Str uc t ur e _

I □Propulsion

■ Rower

□st ructure & Materials

□ Thermal

■ Guidance, Navigation &
Cont rol

□ Communications

^Information & Data
Handling

□ others (Bus)

^Instrumentations
il Communications

□ information & Data
Handling

□r

Reference: Worden.P, Space Warfare Center USA, Clementine 2, ISU August 1996.

Appendix D • 323

Cost matrix of CLUSTER

SOLAR PROBE MISSION:

MISSION COST Break down Sturcture

SPACE SEGMENT COST Break down Structure

□ground segmbmt
■launcher

□ SPACE SEGMB'JT

□ Propuls

□ si ructure & Materials
OThermal

•Guidance, Navigation &
Control

OCommunicat ions

■information & Data
Handling
□ahers (Bus)

■ instrumentations

■Communications

^Information & Data
Handling

□ahers (Payload)

Reference: De Dalmau.J, European Space Agency, personal communications, ISU August 1996.

324 • Ra: The Sun for Science and Humanity

Reference: FAST-Fast Auroral Snapshot Explorer,

http;//sunland.gsfc.nasa.gov/smex/fast/fast Jop.html, NASA.

Appendix D • 325

Cost matrix of SOHO

MISSION COST Break down Structure

SPACE SEGMENT COST Break down Structure

□ Propulsion

■ Power

□ Structure & Materials
o Thermal

■ Guidance, Navigation &
Control

□ Communications

■ Information & Data
Handling

14 % Q Others (Bus)

■ Instrumentations
Cl Communications

□ Information & Data
Handling

□ Others (Payload)

Reference: Wnuk.G, European Space Agency, personal communications, ISU August 1996.

22 %

Cost matrix of TIMED

SOLAR PROBE MISSION;

I TIMED

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CftOCN'D

TRACKING

- n ST A -

■

Taurus da ss( LEO)

S 28.000,000

£>PACfc

"

— -

Propulsior

Power

Structure <Sr

Guidance, Navigation &

Communications

Information & Data

Others (Bus

PAYLOAD

1 1 2Kg, 1 50vv, 6 00 Km circular with 74 4 degree in din a ten

Instrumentations

Communications

s-Ba nd, STD.N compa lihle

»

N Others (Payloadji

♦

6 Li'.' 1 VO pi.-’l

Total exc. Prop el I am

|',1

incl.Propellant/lftoJ

TOTAL i

■ 1

MISSION COST Break down Sturcture

□ GROUNDSEGMENT
■ LAUNCHER

□ SPACE SEGMENT

SPACE SEGMENT COST Break down Structure

□ Propulsion
■ Power

□ Structure & Materials

□ Thermal

■ Guidance, Navigation &
Control

□ Communications

■ information & Data
Handling

□ Others (Bus)

■ Instrumentations

□ Communications

□ Information & Data
Handling

. . . .. _ ... .. _ □ Othprc /Pavlnarn

Reference: TIMED-Thermosphere Ionosphere Mesosphere Energetics and Dynamics, http://sd-

wzuzv.jhuapI.edu/TIMED/ovennezv.html, NASA.

Appendix D • 327

Cost matrix of FIRE Mission, Solar probe

MISSION COST Break down Structure

Reference: Randolph J, JET Propulsion Laboratory USA, FIRE mission, ISU August 1996 .

328 • Ra: The Sun for Science and Humanity

Cost matrix of FIRE Mission, Plamva

MISSION COST Break down Sturcture

Reference: Randolph.J, JET Propulsion Laboratory USA, FIRE mission, JsU Augi 1st 1996

Appendix D • 329

Cost matrix of FIRE Mission, Plamya and Solar probe

SOLAR PROBE MISSION:

MISSION COST Break down Sturcture

SPACE SEGMENT COST Break down Structure

□ propulsic

□ Structure & Materials

__ nThermal

®Guidance, Navigation &

f ) Control

□ Communications

l Blnformation & Data

□Others

10/0 ®Communications

□ information & Data
Handling

□ others (Ravload)_

Reference: Randolph.J, JET Propulsion Laboratory USA, FIRE mission, ISU August 1996.

330 • Ra: The Sun for Science and Humanity

Cost matrix of RA application mission

Cost matrix of SAUNA

■ Communicai ions

110/o □hfotmat ion & Dal a Handfinq !

4%

□ ahere (payload)

Reference: French.L, JET propulsion Laboratory LISA, personal communications, ISU August
1996.

332 • Ra: The Sun for Science and Humanity

Appendix E

Existing and Proposed
Early Warning Systems

This appendix lists some of the solar environment monitoring and early warning agencies
which operate today and provide forecasting and alert services. In section E.2 proposed
systems are described.

E.l Existing Space Environmental Forecasting Services

U.S Systems

The Space Environment Center (SEC) which is part of the National Oceanographic and
Atmospheric Administration (NOAA) in co-operation with the 50th Weather Squadron (50th
WS) of the United States Air Force (USAF) provide a number of space environmental
products to customers in the United States. The two agencies share resources and divide
their customer base in order to serve both military and civilian users which include the
NASA, the Federal Aviation Administration (FAA), HF and SF radio operators, power
companies, satellite operators, radar user, researchers and many more.

SEC and 50th WS use a number of observations in order to issue warnings concerning a
geophysical event, provide short and long term forecast information of space environmental
conditions and to provide anomaly analysis to determine whether a problem experienced by
a customer was related to the space environment. Data is acquired from:

• 2 GOES (Geostationary Observational Environmental Satellite) vehicles
operated by NOAA measuring X-rays, charged particle flux and magnetic field
flux at Geosynchronous altitude.

• 2 DMSP (Defense Military Satellite Program) vehicles operated by the USAF
measure precipitating particle and plasma flux (which give information about
the aurora) along with magnetic variations in low earth polar orbits.

• WIND (Weather Information Display System) operated by NASA provides two
hours of real-time solar wind measurements per day.

• Various other military satellites provide magnetic field flux and particle fluxes
in a variety of orbits.

• The Solar Electro-Optical Network (SEON) operated by the USAF employs five
SOON (Solar Observation Optical Network) telescopes along with four RSTN
(Radio Solar Telescope Network) telescopes providing continuous solar data
from six locations world-wide. SOON provides information concerning the
photosphere and active surface regions (white light images), the chromosphere
(Hydrogen alpha line) and the corona (Calcium K line) while RSTN measure
the Sun's output at a variety of radio frequencies.

• Numerous magnetomers operated by the United States Geological Survey
provide data on the Earth's magnetic field at the surface.

• Various Ionospheric Measuring Systems (IMS) determine the height of the
various layers of the ionosphere and measure total electron content (TEC).

• A riometer provides ionospheric absorption level information at the poles.

• A Neutron monitor measures high-energy particle fluxes at the surface.

Data is collected from the various sources and fed into models that generate warnings, alerts
and forecast information. The Magnetospheric Specification Model (MSM) designed at Rice
University provides data on the magnetosphere while other models used provide data
concerning other regions.

Australian Systems

The IPS Warning Centre in Australia has optical and radio observatories in Culgoora (near
Narrabri, NSW), and Learmonth (near Exmouth, WA). The Learmonth observatory is jointly
operated with the United States Air Force.

Canadian Systems

The Geomagnetic Laboratory, a division of the Geological Survey of Canada, provides
geomagnetic storm alerts and forecasts to Hydro-Quebec's Transmission Control Centre.

E.2 Proposed Solar Threat Monitoring & Early Warning Systems

Under study as of this writing by NASA, the US Air Force, and the University of
Birmingham (U.K.) is a science and applications mission which will have considerable impact
on space weather forecasting systems. The Solar Mass Ejection Imager (SMEI) will image the
Thompson - scattered white light from dense structures in the interplanetary solar wind.
This method is described in Section 9.3.4, and is expected to enable SMEI to image and track
solar mass ejections (CME) and determine with 1 to 3 days notice when one will impact the
Earth. SMEI will image the inner interplanetary region from Earth orbit every 90 minutes and
is hoped to be launched before the next solar maximum in 2001. While this mission could be
extremely valuable for both scientific and applications interests, it is worthwhile noting that
using remote sensing it will not measure the interplanetary magnetic fields destined for
Earth and thus does not satisfy the requirements derived and outlined in Section 9.2.

334 • Ra: The Sun for Science and Humanity

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